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LlBKAKr OP cWgbt.^ 


SUMMARY TECHNICAL REPORT 
OF THE 

NATIONAL DEFENSE RESEARCH COMMITTEE 


DECI . ;\S.3.TPTKD 
■by authority occvoini'y of 

G l» 1 4 ^ i960 

Defense mono 2 A 1960 
LIBRARY OF CONGRESS 


This document contains information affecting the national defense of the 
United States within the meaning of the Espionage Act, 50 U. S. C., 31 and 
32, as amended. Its transmission or the revelation of its contents in any 
manner to an unauthorized person is prohibited by law. 

This volume is classified SECRET in accordance with security regulations 
of the War and Navy Departments because certain chapters contain mate¬ 
rial which was SECRET at the date of printing. Other chapters may have 
had a lower classification or none. The reader is advised to consult the War 
and Navy agencies listed on the reverse of this page for the current classifi¬ 
cation of any material. 


Unclassified 




Manuscript and illustrations for this volume were prepared for 
publication by the Summary Reports Group of the Columbia 
University Division of War Research under contract OEMsr-1131 
with the Office of Scientific Research and Development. This vol¬ 
ume was printed and bound by the Columbia University Press. 

Distribution of the Summary Technical Report of NDRC has 
been made by the War and Navy Departments. Inquiries concern¬ 
ing the availability and distribution of the Summary Technical 
Report volumes and microfilmed and other reference material 
should be addressed to the War Department Library, Room 1A- 
522, The Pentagon, Washington 25, D. C., or to the Office of 
Naval Research, Navy Department, Attention: Reports and Doc¬ 
uments Section, Washington 25, D. C. 


Copy No. 


197 


This volume, like the seventy others of the Summary Technical 
Report of NDRC, has been written, edited, and printed under 
great pressure. Inevitably there are errors which have slipped past 
Division readers and proofreaders. There may be errors of fact not 
known at time of printing. The author has not been able to follow 
through his writing to the final page proof. 

Please report errors to: 


JOINT RESEARCH AND DEVELOPMENT BOARD 
PROGRAMS DIVISION (STR ERRATA) 
WASHINGTON 25, D. C. 


A master errata sheet will be compiled from these reports and sent 
to recipients of the volume. Your help will make this book more 
useful to other readers and will be of great value in preparing any 
revisions. 




SUMMARY TECHNICAL REPORT OF DIVISION 5, NDRC 


VOLUME 1 


GUIDED MISSILES 
AND TECHNIQUES 

DEC LASSIFIED 
Dy authority Sucretury of 

OCi 2 u iHbj 

Defense memo 2 August I960 

LIBRARY OF CONGRESS 

f 

OFFICE OF SCIENTIFIC RESEARCH AND DEVELOPMENT 
VANNEVAR BUSH, DIRECTOR'* 

NATIONAL DEFENSE RESEARCH COMMITTEE 
JAMES B. CONANT, CHAIRMAN 

DIVISION 5 

HUGH H. SPENCER, CHIEF 


WASHINGTON, D.C., 1946 


Unclassified 



NATIONAL DEFENSE RESEARCH COMMITTEE 


James B. Conant, Chairman 
Richard C. Tolman, Vice Chairman 
Roger Adams Army Representative 1 

Frank B. Jewett Navy Representative 2 

Karl T. Compton Commissioner of Patents 3 

Irvin Stewart, Executive Secretary 


1 Army representatives in order of service: 


Maj. Gen. 
Maj. Gen. 
Maj. Gen. 
Brig. Gen. 


G. V. Strong 
R. C. Moore 
C. C. Williams 
W. A. Wood, Jr. 


Col. L. A. Denson 
Col. P. R. Faymonville 
Brig. Gen. E. A. Regnier 
Col. M. M. Irvine 


Col. E. A. Routheau 


2 Navy representatives in order of service: 

Rear Adm. H. G. Bowen Rear Adm. J. A. Furer 

Capt. Lybrand P. Smith Rear Adm. A. H. Van Keuren 

Commodore H. A. Schade 
. 3 Commissioners of Patents in order of service: 

Conway P. Coe Casper W. Ooms 


NOTES ON THE ORGANIZATION OF NDRC 


The duties of the National Defense Research Committee 
were (1) to recommend to the Director of OSRD suitable 
projects and research programs on the instrumentalities of 
warfare, together with contract facilities for carrying out 
these projects and programs, and (2) to administer the tech¬ 
nical and scientific work of the contracts. More specifically, 
NDRC functioned by initiating research projects on requests 
from the Army or the Navy, or on requests from an allied 
government transmitted through the Liaison Office of OSRD, 
or on its own considered initiative as a result of the experi¬ 
ence of its members. Proposals prepared by the Division, 
Panel, or Committee for research contracts for performance 
of the work involved in such projects were first reviewed by 
NDRC, and if approved, recommended to the Director of 
OSRD. Upon approval of a proposal by the Director, a con¬ 
tract permitting maximum flexibility of scientific effort was 
arranged. The business aspects of the contract, including 
such matters as materials, clearances, vouchers, patents, 
priorities, legal matters, and administration of patent matters 
were handled by the Executive Secretary of OSRD. 

Originally NDRC administered its work through five 
divisions, each headed by one of the NDRC members. 
These were: 

Division A—Armor and Ordnance 
Division B—Bombs, Fuels, Gases, & Chemical Problems 
Division C—Communication and Transportation 
Division D—Detection, Controls, and Instruments 
Division E—Patents and Inventions • 


In a reorganization in the fall of 1942, twenty-three ad¬ 
ministrative divisions, panels, or committees were created, 
each with a chief selected on the basis of his outstanding 
work in the particular field. The NDRC members then be¬ 
came a reviewing and advisory group to the Director of 
OSRD. The final organization was as follows: 

Division 1—Ballistic Research 

Division 2—Effects of Impact and Explosion 

Division 3—Rocket Ordnance 

Division 4—Ordnance Accessories 

Division 5—New Missiles 

Division 6—Sub-Surface Warfare 

Division 7—Fire Control 

Division 8—Explosives 

Division 9—Chemistry 

Division 10—Absorbents and Aerosols 

Division 11—Chemical Engineering 

Division 12—Transportation 

Division 13—Electrical Communication 

Division 14—Radar 

Division 15—Radio Coordination 

Division 16—Optics and Camouflage 

Division 17—Physics 

Division 18—War Metallurgy 

Division 19—Miscellaneous 

Applied Mathematics Panel 

Applied Psychology Panel 

Committee on Propagation 

Tropical Deterioration Administrative Committee 


iv 


Library of Congress 




2015 


490932 


































NDRC FOREWORD 


as events of the years preceding 1940 revealed 
more and more clearly the seriousness of the 
world situation, many scientists in this country came 
to realize the need of organizing scientific research for 
service in a national emergency. Recommendations 
which they made to the White House were given care¬ 
ful and sympathetic attention, and as a result the 
National Defense Research Committee [NDRC] was 
formed by Executive Order of the President in the 
summer of 1940. The members of NDRC, appointed 
by the President, were instructed to supplement the 
work of the Army and the Navy in the development 
of the instrumentalities of war. A year later, upon 
the establishment of the Office of Scientific Research 
and Development [OSRD], NDRC became one of its 
units. 

The Summary Technical Report of NDRC is a 
conscientious effort on the part of NDRC to sum¬ 
marize and evaluate its work and to present it in a 
useful and permanent form. It comprises some seventy 
volumes broken into groups corresponding to the 
NDRC Divisions, Panels, and Committees. 

The Summary Technical Report of each Division, 
Panel, or Committee is an integral survey of the work 
of that group. The first volume of each group’s report 
contains a summary of the report, stating the prob¬ 
lems presented and the philosophy of attacking them, 
and summarizing the results of the research, develop¬ 
ment, and training activities undertaken. Some vol¬ 
umes may be “state of the art” treatises covering 
subjects to which various research groups have con¬ 
tributed information. Others may contain descrip¬ 
tions of devices developed in the laboratories. A 
master index of all these divisional, panel, and com¬ 
mittee reports which together constitute the Sum¬ 
mary Technical Report of NDRC is contained in a 
separate volume, which also includes the index of a 
microfilm record of pertinent technical laboratory re¬ 
ports and reference material. 

Some of the NDRC-sponsored researches which 
had been declassified by the end of 1945 were of suf¬ 
ficient popular interest that it was found desirable to 
report them in the form of monographs, such as the 
series on radar by Division 14 and the monograph on 
sampling inspection by the Applied Mathematics 
Panel. Since the material treated in them is not 
duplicated in the Summary Technical Report of 


NDRC, the monographs are an important part of the 
story of these aspects of NDRC research. 

In contrast to the information on radar, which is of 
widespread interest and much of which is released to 
the public, the research on subsurface warfare is 
largely classified and is of general interest to a more 
restricted group. As a consequence, the report of Di¬ 
vision 6 is found almost entirely in its Summary 
Technical Report, which runs to over 20 volumes. 
The extent of the work of a Division cannot therefore 
be judged solely by the number of volumes devoted 
to it in the Summary Technical Report of NDRC; 
account must be taken of the monographs and avail¬ 
able reports published elsewhere. 

Division 5 was responsible for research and devel¬ 
opment work on guided missiles. As the war came to 
an end, work in this field rated a priority second only 
to that of the work on atomic fission, for its implica¬ 
tions for any future war were seen to justify such 
evaluation. Thus, Division 5 both contributed to the 
winning of the war and laid a solid technical founda¬ 
tion for future research aimed at keeping the nation 
prepared against any emergency. 

The Division, under the leadership first of Harold 
B. Richmond and later under Hugh H. Spencer, suc¬ 
ceeded in developing a glide bomb that homed on its 
target by radar, and a visually guided high-angle 
bomb with radio remote control. These weapons were 
employed to destroy key communication links in 
Italy, France, and Burma, and Japanese shipping 
and naval units. The war’s end prevented combat use 
of a heat-homing high-angle bomb, a television-guided 
medium-angle bomb, and a glide bomb promising 
very many times the accuracy of a conventional 
bomb. 

The work of Division 5 is described in this Sum¬ 
mary Technical Report. Preparation of the report 
was supervised by the Division Chief, and its publica¬ 
tion has been authorized by him. To him and his col¬ 
leagues, for helping to keep our armed forces in the 
forefront of the technical race which is modern war¬ 
fare, we express our sincere appreciation. 

Vannevar Bush, Director 
Office of Scientific Research and Development 
J. B. Conant, Chairman 
National Defense Research Committee 



v 
































































































































-* 

































FOREWORD 


T his summary technical report presents the 
guided-missile program carried out by the Na¬ 
tional Defense Research Committee [NDRC] during 
the war. The sponsor, Division 5, was created by the 
reorganization of NDRC some eighteen months after 
the start of NDRC’s program of war research. Some 
of the work here reported, therefore, treats of projects 
completed by predecessor groups. Some represents 
activities which had been begun under the super¬ 
vision of predecessors and were completed by Di¬ 
vision 5. 

The technical information presented here stems 
from the work of the Division’s contractors. This 
principle is not unique with Division 5 but its reitera¬ 
tion is perhaps worth while. The function of the 
Division was one of critical administration rather 
than the exploration of new fields of scientific thought 
or the development of new techniques in applied 
science. It considered possible new fields of scientific 
study and assigned to contractors specific investiga¬ 
tions within those fields. In addition, it directed the 
activities of contractors in the technologies necessary 
to bring to fruition the results of its contractors’ 
scientific research. 

Having established a group of working contracts, 
the Division then critically reviewed the work of the 
contractors and interpreted their results to the Serv¬ 
ices. It stimulated where additional efforts seemed in¬ 
dicated, and restrained where diversification of effort 
seemed profitless. 

The material presented here, therefore, is a sum¬ 
mary of the developments and discoveries of the 
Division’s contractors. Credit for whatever is new 
herein is due contractors and not the Division, and 
where the text does not make this clear an oversight 
has occurred. The report is, however, something more 
than a mere distillate of contractors’ reports. A large 
group of competent scientific and technological men 
and women was assembled by the Division’s con¬ 
tractors, all working in the broad field of guided mis¬ 
siles. The views of these individuals were often di¬ 
vergent, occasionally contradictory. The book at¬ 
tempts to bring these divergencies into focus and to 
resolve them, to draw attention to contradictions and 
to assess the arguments supporting the hypotheses 
set forth. 

The book is divided into four parts: a summary, 
followed by the body of the report in three parts. 


The summary presents a resume of the entire ac¬ 
tivities of the Division. It is intended to be sufficient 
to meet the needs of those generals and admirals 
having responsibilities in the guided-missile field but 
who are prevented by the pressure of other duties 
from undertaking its detailed study. 

The Division directed the development of four 
systems of guided missiles. Two of these systems in¬ 
volved remote radio control; two were automatically 
target-seeking. Part I discusses these systems in de¬ 
tail. The whole experience of the Division proves 
conclusively that guided-missile development can be 
successfully prosecuted only by careful consideration 
of the integrated missile system—airframe, control 
surfaces, and means of guidance. This principle is 
fully supported by what has been reported of the ex¬ 
perience of our enemies. Two of the missile systems 
developed by the Division reached combat. Another 
was ordered for combat, and combat teams were in 
training. A fourth was still incompletely developed 
as hostilities ceased. These four systems are broadly 
considered in Part I. 

Part II discusses the components which the Divi¬ 
sion undertook to develop separately, as distin¬ 
guished from the development of a system consisting 
of a missile and its controls. Only a few of the many 
proposed were prosecuted, the experience of the Di¬ 
vision clearly teaching that such separation of effort 
can almost never lead to success. In addition to a 
description of the separate projects in control of mis¬ 
siles, Part II describes accessory techniques, such as 
the development of simulators and trainers, which 
the Division found to be of important assistance to 
its program. 

Part III presents direct contributions from certain 
of the Division’s contractors. Each of the groups 
having primary responsibility for a guided-missile 
system was invited to contribute a monograph chap¬ 
ter to this report. Dr. Hugh L. Dryden of the Na¬ 
tional Bureau of Standards, R. D. Wyckoff of the 
Gulf Research and Development Company, and Dr. 
W. B. Klemperer of the Douglas Aircraft Company 
accepted the invitation; the Massachusetts Institute 
of Technology declined. 

Each of these authors is peculiarly qualified to 
write critically in the guided-missile field. Besides his 
work for the Division as director of the glide-bomb 
program, Dr. Dryden has served on the von Karman 


i vii 

Uncl ^ifi e! i 


FOREWORD 


viii 


Committee of the Chief of Air Staff. This committee 
conducted a very broad investigation into the future 
possibilities of aerial combat. Mr. Wyckoff directed 
the development program of the Division’s high- 
angle dirigible bomb, the only guided missile save the 
Japanese suicide devices to see combat use in signifi¬ 
cant quantity during the war. Dr. Klemperer di¬ 
rected the development of Project Roc at Douglas 
Aircraft Company. He brought to the project ex¬ 
perience in both the airship and airplane fields. His 
knowledge gained from experience in the Roc project 
has been enriched by an extended tour of Germany 
since the cessation of hostilities there. As a member 
of the ALSOS Mission (under G-2, U. S. Army In¬ 
telligence) he was charged with the responsibility of 
evaluating the wartime development resources of 
Germany in the guided-missile field. 

The facts presented in Parts I and II are the con¬ 
tractors’; the comments thereon are the authors’. In 
Part III the facts and their evaluation are the sole 
product of the contractors. No effort has been made 
to resolve differences which may appear between the 
points of view expressed in these chapters and those 
in the chapters of Parts I and II. Progress in science 
and in technology is made by the candid recognition 
of divergent views and by their resolution through 
frank discussion. 

The book as a whole is repetitious. This is neither 
good nor inadvertent. It is the price paid for making 
each chapter self-sufficient. Thus the reader whose 
prime interest is in glide bombs need not read the 
chapter on Felix, although many problems are com¬ 
mon to both projects and the findings in one program 
are apposite to the other. 

Mathematics has been avoided but not shunned. 
The use of this language has been invoked when that 
form of speech (1) by its succinctness avoids verbiage; 
(2) by its exact limitation avoids the danger of gen¬ 
eralization not justified by the experimental evi¬ 
dence; and (3) by its generality concisely states the 


scope of a conclusion. The editor has attempted to 
prepare the text so that the program will be clear 
even to those whose linguistics do not include mathe¬ 
matics; in general, therefore, neither the quantity of 
the mathematics nor its level is such as to repel the 
reader who has a basic technical background. 

The book makes no attempt to recommend a 
broad program on guided missiles for the United 
States. Many other groups, both military and civil¬ 
ian, have undertaken that responsibility. What is at¬ 
tempted here is the presentation of such experience 
of the Division and the formulation of such principles 
arising from it as may profitably be brought to bear 
on the guided-missile program of this country as it 
develops. 

In writing, the editor has been trapped into use of 
the phrase, “So far as the Division is aware . . . ” 
This introduction to each statement is accurate but 
dishonest. As a result of considerable demobilization, 
the Division now consists of the editor and two tech¬ 
nical aides with whom he is in continual touch. Seri¬ 
ous effort has been made to obtain from former Divi¬ 
sion members their criticisms of the material as it 
has appeared. They have been most generous with 
their time and thought. The book should not, how¬ 
ever, be taken as the outcome of deliberations by the 
men who formerly comprised the body of the Divi¬ 
sion. Where errors of omission, of emphasis, or of 
fact occur, the fault is the editor’s. 

A Summary Technical Report is hardly the me¬ 
dium for the expression of appreciation to colleagues. 
The editor can hardly close this foreword, however, 
without such a word of gratitude to all his former 
associates, who are enumerated in the list of OSRD 
appointees. Particularly is he indebted to Dr. E. 
W. Phelan and Dr. J. C. Boyce for their direct 
contributions to the text. 

Hugh H. Spencer 
Chief, Division 5 



CONTENTS 


CHAPTER PAGE 

Summary. 1 

PART I 

GUIDED-MISSILE SYSTEMS 

1 Glide Bombs by Hugh H. Spencer . 7 

2 Azon and Razon by Hugh H. Spencer .... 27 

3 Felix and Dove Eye by Hugh H. Spencer ... 48 

4 Roc and its Controls by Hugh H. Spencer ... 73 

PART II 

COMPONENTS AND ACCESSORY ACTIVITIES 

5 Television by Hugh H. Spencer .93 

6 Radio Control Systems by Hugh H. Spencer . . 134 

7 Servomechanisms by Joseph C. Boyce . . . . 146 

8 Instrumentation by Hugh H. Spencer . . . . 157 

9 Miscellaneous Control Systems by Hugh H. Spencer 167 

10 Simulation by Earl W. Phelan .202 

11 Transition and Engineering Activities by Hugh H. 

Spencer .223 

PART III 

CONTRIBUTIONS FROM INDIVIDUAL 
INVESTIGATORS 

12 Some Aspects of the Design of Homing Aero- 

Missiles by Hugh L. Dry den .231 

13 The Design of High-Angle Dirigible Bombs by 

R. D. Wyckoff . ".257 

14 Project Roc in Retrospect by W. B. Klemperer . 266 

APPENDIX 

A Interception and Escape Techniques at High 
Speed and Altitude by W. P. Klemperer . . . 285 

B Lateral Stability of Homing Glide Bombs 
with Application by Navy SWOD Mark 7 and 

Mark 9 by H. K. Skramstad .335 

C Infrared Radiation and Its Application to 
Homing Missiles by Alan C. Bemis .348 









CONTENTS 


Glossary.359 

Bibliography.363 

OSRD Appointees.368 

Contract Numbers.369 

Service Project Numbers.371 

Index.373 



























. 


























































































■ 















































































. 





















Azon strikes on Taungup Road Bridge in Burma by Tenth Air Force, June 1945. Altitude, 12,000 feet, 



SUMMARY 


T he successful glide-bomb attack by the Ger¬ 
mans on a British convoy in the Bay of Biscay 
on August 23, 1943, gave an impetus to American 
development of guided missiles which has intensified 
continuously ever since. This is not to say that Amer¬ 
ican effort had not been concerned with weapons of 
this character before the German guided bombs made 
their appearance. On the contrary, the United States 
Services had been working for some years on guided 
bombs, both powered (drones) and unpowered (glide 
bombs). In addition, NDRC had had guided-missile 
projects under development for many months. The 
one successful combat test by the Germans of Hs 293 
in August 1943, followed by their use of FX-1400 
against the USS Savannah a few weeks later, how¬ 
ever, did more to bring to the attention of senior 
executives at the planning level the possibilities in 
guided missiles than all the previous demonstrations 
and reports by research and development groups. 
This was perhaps as true in the Services as it was 
within NDRC. The V-l buzz-bomb and the long- 
range V-2 rocket, while not guided in the sense that 
their trajectories could be altered after launching, 
stimulated a continuing interest. As World War II 
closed, there was probably no military development 
program, with the exception of atomic power, of 
higher priority than guided missiles. 

NDRC had four systems of guided missiles in its 
development program. Two of these saw combat use, 
and a thiird was in production for combat with crews 
in train ng at the end of hostilities. These systems 
comprised radar-homing glide bombs, visually guided 
high-angle bombs with radio remote control, a heat¬ 
homing high-angle bomb, and a medium-angle bomb 
of high maneuverability guided by television. In ad¬ 
dition, the program included the development of a 
few of the many components which were suggested 
to control or to assist in the control of bombs. 

Glide Bombs 

The glide bombs developed by Division 5, NDRC, 
were high-wing monoplanes which were character¬ 
ized by two features. They had a high-wing loading 
so that they were self-supporting only at speeds con¬ 
siderably greater than those of conventional bom¬ 
bardment aircraft; they consequently had to be car¬ 
ried under the airplanes that attacked with them. 


Unlike the glide bombs of the Army Air Forces’ 
program, they had no rudders or elevators. Con¬ 
trolled flight was obtained by trailing-edge wing flaps 
which altered the lift developed by the wings. These 
flaps were controlled through a linkage which per¬ 
mitted the flaps to be raised and lowered either ex¬ 
actly together or in opposition—one up, the other 
equally down. Simultaneous elevation or depression 
of the flaps provided control in range. Differential 
operation of the flaps provided control in azimuth. 
This design produced in the missile the characteristic 
of nearly constant angle between the fuselage and 
the line of flight for all inclinations of the glide path. 

Thus a homing device mounted in the glide bomb 
could be bore-sighted so that its axis of scan was al¬ 
ways approximately tangent to the flight path. All 
the work of the Division has indicated that this is 
the most important property of homing missiles. Un¬ 
less the homing device scans along the line of flight, 
a computing device has to be inserted in the servo 
system to correct continuously for angular deviations 
between axis of scan and the flight path. The homing 
missiles of the Division’s program were designed to 
look where they were going. 

The radar-homing control equipment for the glide 
bombs came in two versions. The initial version, 
Pelican, was a Division project. The radar trans¬ 
mitter which illuminated the target was mounted on 
the carrying aircraft. The glide bomb carried the re¬ 
ceiver. Some restriction was thus placed on the 
maneuvers which the aircraft could make after re¬ 
lease, since it had to maintain a position and attitude 
which would permit continuous illumination of the 
target until impact. Actually, this restriction was not 
serious, as the transmitting equipment (ASG) had a 
considerable range and a substantial field. A very 
satisfactory maneuver was to drop the missile at a 
range of 12 to 15 miles, make a 180-degree turn, 
and withdraw. Such a maneuver kept the attacking 
aircraft beyond the range of antiaircraft defense while 
still providing strong radar illumination throughout 
the 4-minute flight of the missile. 

In the second version, Bat, in which the Division 
acted as consultants to the Navy Bureau of Ordnance 
on the radar while continuing in a position of prime 
responsibility for the entire system, both the radar 
transmitter and receiver were mounted on the mis¬ 
sile. This design removed all restrictions on the 


1 


2 


SUMMARY 


maneuverability of the aircraft after release, but the 
operation of the fourth-power law in the relationship 
between received signal and range to the target in¬ 
troduced a problem in automatic gain control that 
was solved only by restricting the launching range 
to a maximum of some 7 miles. The range restriction 
with Bat was deemed by the Navy to be of less 
severity than the maneuverability restriction with 
Pelican. Accordingly, Bat was pushed to completion 
and was successfully used against Japanese shipping 
and naval units during the last months of the war. 

The project is continuing under naval direction, 
working toward the addition of lead prediction so 
that the missile need not fly a pursuit course against 
a moving target or a stationary target in the presence 
of wind. 

A television-guided version, Robin, was success¬ 
fully demonstrated early in 1943 but was not prose¬ 
cuted because of its parallel development with the 
AAF glide bomb GB-4. 

Azon and Razon 

Two versions of controllable high-angle bombs 
were developed. Each was visually sighted and re¬ 
motely controlled by radio. The objective was to 
develop a missile of considerably greater precision 
than the standard high-level bomb, but which could 
be carried in and released from the existing racks in 
the bomb-bays of standard aircraft. The main ob¬ 
jective, therefore, precluded the use of any support¬ 
ing aerodynamic surfaces much larger than the tail 
fin of a standard bomb. Aerodynamic control had to 
be obtained, therefore, from the bomb body itself. 

Elimination of roll was important since in control 
of the bomb the identity of rudders and elevators— 
azimuth control and range control—had to be pre¬ 
served. The problem was solved by ailerons. The 
ailerons were controlled by a free gyro (which pre¬ 
served the bomb’s orientation) and a rate gyro (which 
damped out roll oscillation). This problem is more 
acute in a bomb which is controlled in both senses, 
range and azimuth, than in one that is controlled in 
one sense only. With a conventional tail structure 
the simultaneous application of control in yaw and 
pitch produces roll torques which are more difficult 
to cope with than are the roll torques due to the un¬ 
intentional asymmetries usually present in produc¬ 
tion bombs. 

The Division, therefore, pressed the development 
of Azon, a visually guided bomb remotely controlled 


in azimuth only. Its successful demonstration to the 
military at Muroc, California, occurred on Septem¬ 
ber 10, 1943, a fortnight after the appearance of the 
first German guided missile. A rush production pro¬ 
gram was undertaken, and, in March 1944, a group 
of B-17 aircraft equipped with radio-control trans¬ 
mitters for Azon left for the Mediterranean Theater 
of Operations. The missile was effective there against 
transportation links of the enemy forces which were 
resisting the Fifth Army’s advance in the Italian 
campaign. In particular, the Avisio Viaduct south of 
the Brenner Pass was closed by Azon. Other success¬ 
ful operations with Azon against the locks of the 
Iron Gate on the Danube led to the acceleration of 
Azon production. 

Use of Azon in the European Theater was less suc¬ 
cessful, although successful missions were flown 
against key bridges on the Seine and Loire just prior 
to and during the Normandy operation. The prin¬ 
cipal reason for lack of success in this theater seems 
to have been organizational rather than technical. 
Specifically, the policy of evaluating the efficiency 
of a squadron on the basis of the tonnage of bombs 
dropped rather than on the number of targets de¬ 
stroyed militated seriously against the success of 
Azon in the ETO. However, in the Burma campaign, 
December 1944 to the end of hostilities, Azon was. 
strikingly effective, thoroughly disrupting Japanese 
communications by repeatedly cutting bridges on the 
Taungup Road and the Bangkok-Chiengmai railway. 

The accomplishment of remote control in both co¬ 
ordinate axes, range and azimuth, was more difficult. 
The normal cruciform tail-fin structure was aban¬ 
doned in favor of an octagonal shroud, which is only 
slightly subject to roll torque due to simultaneous 
application of control in yaw and pitch. As a further 
assurance of roll stability, the area size of the ailerons 
and their speed of operation were increased. 

A more important problem than roll stabilization, 
however, was that of parallax. It is impossible for an 
observer in the bombardment aircraft to estimate 
the range error of his bombs accurately since he looks 
along the plane of the bomb trajectory. The Germans 
solved this problem in FX-1400, their high-angle 
dirigible bomb, by a maneuver of the airplane which 
placed it well abaft the missile. This maneuver re¬ 
sults in nearly stalling out the airplane, and was 
deemed by the Division unacceptable for combat. 

Instead, the Division developed with the coopera¬ 
tion of the NDRC Fire-Control Division, the Crab 
attachment to the standard M14 bombsight. This 



SUMMARY 


3 


very simple attachment superposed an image of 
Razon—the high-angle bomb controlled in range and 
azimuth only—onto the image of the terrain as seen 
through the bombsight at the point where the bomb 
would fall with no further control. This development 
permitted the bombardier to correct errors in aiming 
throughout the time of flight of the bomb. 

The improvement in accuracy due to the use of 
Azon compared with standard bombing in some thir¬ 
ty times. That is to say for the same high 'probability 
of obtaining a hit on a target approximately 50 ft 
square, thirty times as many individually aimed 
standard bombs have to be dropped as are required 
with Azon. This figure of merit was developed by the 
NDRC Applied Mathematics Panel based on test 
data taken at the AAF Proving Ground and at the 
evaluation base of the Air Forces Board. 

The improvement with Razon is some three hun¬ 
dred times. The gain due to range control—ten times 
—is less than the gain due to azimuth control since 
errors in estimation of time of flight, which contribute 
to the range error, cannot be corrected by the Crab 
sight. The addition of Jag (Just Another Gadget) 
eliminated about two-thirds of the range error caused 
by error in estimation of the time of flight, resulting 
in an overall improvement of some 700 times in ac¬ 
curacy over conventional bombing. 

Units to take Razon into combat were in training 
when the war closed. 

Felix 

The visually guided high-angle bomb required a 
continuous bomb run from the release point until the 
instant of impact. To eliminate this requirement and 
to provide the possibility of precise night bombard¬ 
ment, the Division developed Felix, a heat-homing 
high-angle bomb. 

The distinction between heat and other infrared 
radiations must be stressed even in this summary. 
For reasons that are fully developed in Chapter 3 
and Appendix C, the Division concentrated on heat 
detectors in the region between 8.5 and 15.0 microns. 
The water-vapor absorption of infrared rays at 
other wavelengths make their use for missile control 
impracticable. 

The statement is repeatedly made that the Ger¬ 
mans had infrared detectors very much more sensi¬ 
tive than we had. So far as heat-homing missiles are 
concerned, this statement is insignificant without 
quantitative reference to the wavelength band. In 


the wavelength band where heat homing is practi¬ 
cable, the statement is not true. 

The aerodynamic structure of Felix was substan¬ 
tially that of the Razon—a standard 1,000-lb GP 
bomb with an octagonal control tail. Stabilization in 
roll was accomplished by an identical gyro system as 
for Razon, controlling similar ailerons. 

The missile was guided to the target by a heat- 
sensitive element which scanned the terrain toward 
which the bomb was falling and directed it toward 
the quadrant—up, down, right, or left—which ra¬ 
diated the greatest heat flux. Thus a thermal target 
such as a steel mill at night attracted Felix by virtue 
of the heat which it radiated. The missile structure 
developed considerable angle of attack in altering its 
trajectory. In order to make the axis of scan lie ap¬ 
proximately tangent to the line of flight, the scan¬ 
ning head was mounted in gimbals and connected to 
the rudders and elevators. This is an approximation 
to the requirement that a homing missile look where 
it is going. A precise fulfillment of this requirement 
involves transient analysis of the response of the mis¬ 
sile as an aerodynamic body. Techniques for such an 
analysis do not as yet exist. 

The sensitivity of the heat-scanning head—10~ 7 
watt per sq cm—was sufficient to make Felix an ef¬ 
fective missile against many targets. Much further 
study of target heat radiation is required before the 
full effectiveness of Felix can be assessed and the 
advisability of developing more sensitive'heat detec¬ 
tors determined. In any case, a study of the absorp¬ 
tion of radiation in common atmospheres from the 
visible range out to 20.0 microns is seriously 
needed. 

The missile was successfully tested against Chan¬ 
nel Key, Florida, and units were in preparation for 
combat by the Twentieth Air Force as World War II 
closed. 

Roc 

Roc was a highly maneuverable medium-angle 
bomb developed by the Division as a homing bomb. 
Initial aims to utilize radar homing were frustrated 
when experiments by the Division’s radar group dis¬ 
covered that satisfactory resolution at microwave 
frequencies could not be obtained at the glide angles 
which Roc could attain. The use of television was 
then planned. 

As the war closed, development was incomplete. 
The project was transferred to the Army Air Forces 



4 


SUMMARY 


when the research and developmental activities of 
the Division terminated. 

Components and Techniques 

In addition to the four missile systems just men¬ 
tioned, the Division prosecuted development pro¬ 
grams on various component devices and accessory 
techniques. Some of them are sufficiently noteworthy 
to merit review in this summary. In particular, a 
general conclusion which the experience of the Divi¬ 
sion emphatically taught should be pointed out. The 
independent development of components without re¬ 
gard to the dynamic characteristics of all the components 
which comprise a guided-missile system leads in 
general to failure. Overall responsibility for the com¬ 
plete system must be placed in a single coordinated 
group. 

At present, the development of the guided-missile 
art does not permit the evaluation of the dynamic 
characteristics of the elements—for example, homing 
device, servomechanism, and airframe—which com¬ 
prise a guided-missile system. Indeed no suitable vo¬ 
cabulary has been formulated for expressing them. 
Broad research studies involving such subjects 
as transient aerodynamic responses are seriously 
needed. 

Television 

The Division studied the application of television 
to guided missiles. Compact, lightweight television 
equipment utilizing carrier frequencies of 100, 300, 
800, and 1,800 me was developed. Amplitude modu¬ 
lation and frequency modulation were explored. In 
general, it can be said that no television-missile sys¬ 
tem was successful as World War II closed, although 
spectacular improvement in pickup sensitivity had 
been accomplished through the development of the 
image orthicon camera tube. 

A problem, unsolved at the end of hostilities, was 
the interference between two television signals re¬ 


ceived at the controlling plane. Both of these signals 
originated at the television bomb: one was trans¬ 
mitted to the aircraft directly, the other was reflected 
to the aircraft from the surface of the ground. These 
two paths were continuously changing in length, the 
direct path lengthening as the bomb approached the 
ground and the ground-reflected path decreasing. 
Thus the two signals were received at apparently 
different frequencies, and their interference produced 
a moving pattern of bars across the received picture. 
The spacing, orientation, and motion of the bars in 
the pattern depended upon the relative velocity of 
the airplane and the bomb. 

The only cure appears to be to use a television 
transmitting antenna so directional that no signal 
can reach the ground to be reflected upward to cause 
interference. In order to get directional antennas of 
a reasonable size and low aerodynamic drag it was 
necessary to go to higher frequencies. Work on a 
television transmitter for Roc operating at a carrier 
frequency of 1,800 me was in progress as hostilities 
ceased. 

Simulation 

The problem of guiding a missile is one of compli¬ 
cated dynamics. The differential equations probably 
are at least of higher order than the second; the co¬ 
efficients are, in general, nonlinear. Simulation is the 
only means readily at hand for coping with such 
problems. 

The Division made a start in the development of 
the art of simulative solution of the motion of a 
guided missile. Much more work needs to be done. 
Especially required is a method of assessing a par¬ 
ticular problem to determine whether the most eco¬ 
nomical attack lies through conventional mathema¬ 
tics with the necessary broad assumptions, through 
point-by-point analysis, through the use of some of 
the modern computing assemblies such as the Rocke¬ 
feller differential analyzer, or through the design of 
a special simulative device. 



PART 1 

GUIDED-MISSILE SYSTEMS 



Chapter 1 


GLIDE BOMBS 


INTRODUCTION 

T wo broad purposes underlie the development 
of guided missiles: the improvement in accuracy 
and the extension of range so that the release point 
for the missile lies beyond the range of lethal anti¬ 
aircraft defense. The Division attacked both prob¬ 
lems simultaneously in the glide-bomb development 
(the Washington Project). Through the addition of 
wings to give increase in lift the range was extended. 
The development of a homing system—either manu¬ 
ally through remote control and television or with a 
wholly automatic radar system—greatly increased 
the accuracy. 

Even before the outset of the glide-bomb project 
in NDRC, remotely controlled aircraft had been 
flown as gunnery targets. A glide bomb was under 
development by the Army Air Forces and, in a pre¬ 
set unguided version, was ready for standardization 
as a combat weapon. In the Navy similar work was 
going forward on pilotless, engine-driven aircraft and 
gliders 3 remotely guided into a target. 

In order to make the axis of received intelligence 
—i.e., television or radar—continuously tangent to 
the flight path, the missile was designed to fly with 
an angle of attack as nearly constant as possible. To 
achieve maximum range from the release point to 
the target the missile was made aerodynamically 
clean. 

These principles were embodied in the three mis¬ 
siles shown in Figure 1: Robin—a missile of 12-ft 
wing span, carrying a 2,000-lb GP bomb and guided 
by television; Pelican—a missile of 8-ft wing span, 
carrying a 325-lb depth charge and self-guided by 
radar reflections from signals radiated by a trans¬ 
mitter located in the carrying aircraft; Bat—a mis¬ 
sile of 10-ft wing span, carrying a 1,000-lb GP bomb 
and self-guided by radar reflections from signals ra- 


a The distinction between gliders and glide bombs is a subtle 
one; of course, strictly speaking, the latter forms a specific 
group of the former. In this volume each term will carry a 
specific implication. Gliders are nonpowered winged missiles 
whose wing loading is low enough to make them self-support¬ 
ing at the flying speed of ordinary aircraft; they can, in conse¬ 
quence, be towed. Glide bombs are nonpowered, winged missiles 
whose heavy wing loading requires a flying speed above that of 
conventional aircraft; they must, therefore, be carried. 


diated by a transmitter located in the missile itself. 
This combined transmitter-receiver equipment was 
developed for the Navy, the Division carrying the 
responsibilities only of a consultant. 

Of these missiles one (Bat) reached combat. Robin 
reached a point of sufficient development to achieve 
the accuracy inherent in existing television. The rela¬ 
tively flat glide angle produces considerable fore¬ 
shortening of the television picture, which makes it 
extremely difficult for the controlling bombardier to 
estimate and to correct the error in the range sense. 
Even with the improved quality of television now 
available, the desirability of pursuing the develop¬ 
ment of television-guided glide bombs is doubtful. 

By the time Pelican was sufficiently developed for 
combat use, the submarine menace had been largely 
eliminated by other methods; a larger payload was 
required for other targets, and the limitation in 
maneuver placed on the airplane carrying the radar 
transmitter made the prosecution of the develop¬ 
ment of Bat in the 1,000-lb size more attractive to 
the using Service than the development of a 1,000-lb 
Pelican. The Bat method of control, however, poses 
problems not found in Pelican. 

I 2 GENERAL 

The general aspects of homing-bomb design are 
treated in some detail in Chapter 12. Barely more 
than a recitation of the problems which have to be 
solved will be attempted here. 

121 Coordination of Missile 

and Control 

The problem which has proved of greatest impor¬ 
tance to this Division has not been one of pure 
science nor—in its most limited meaning—one of 
technology. It is what we have come to call systems 
engineering and has, perhaps, been most aptly ex¬ 
pressed by Dry den in Chapter 12: 

“The impression is prevalent that scientific ad¬ 
vances in many fields have progressed to the point 
where the development of such a missile is purely a 
matter of engineering design on the part of specialist 


7 



8 


GLIDE BOMBS 





Figure 1. Washington Project television and radar glide bombs: Robin—television (top); Pelican—RHB radar (center); 
Bat—SRB radar (bottom). 












GENERAL 


9 


groups with the usual coordination as to dimensional 
requirements, weights, and time of completion. Ex¬ 
perience has taught otherwise. Optimistic time sched¬ 
ules based on such an assumption cannot be met. 
The development of successful homing aero-missiles 
requires the solution of certain research problems as¬ 
sociated with the complete article, involving complex 
relationships between the performance characteris¬ 
tics of the component parts. There is required a type 
of overall technical coordination beyond that required in 
the design of aircraft as ordinarily practiced .” (Italics 
ours. Ed.) 

Where this principle has been observed in the 
projects of this Division, a measure of success has 
been achieved (see Chapter 2 particularly); where it 
has been disregarded, failure has always resulted. It 
is noteworthy also that the Japanese, in the develop¬ 
ment of an infrared-homing glide bomb, fell into pre¬ 
cisely this trap; in spite of having competent scien¬ 
tific personnel on each segment of the problem, they 
failed through absence of strong coordination be¬ 
tween the investigating groups. 

122 Target Discrimination and Tracking 

Target discrimination in television depends upon 
the ability of the television equipment to furnish the 
bombardier with a clear picture of the target area. 
This in turn depends upon the sensitivity, resolution, 
and contrast capability of the pickup tube and the 
reliability of the radio link. These problems are dis¬ 
cussed fully in Chapters 5 and 6. 

With automatic radar homing the judgment of a 
human operator is not available for discriminating 
between the multiplicity of echoes which are usually 
returned from the target area. For certain special 
cases, e.g., an isolated aircraft in a cloudless sky, 
no ambiguity exists. In other instances, as with ships 
at sea, discrimination can be obtained by limiting 
the illuminated area to include only the target to be 
attacked. In pulsed radar there is also the possibility 
of utilizing the time of travel from transmitter to 
target to receiver as a means of discrimination. In 
this method it is necessary to have an automatic 
range-tracking element in the circuit so that the 
homing device will continue to respond to echoes 
from a single target as its attack is pushed home. 

Range tracking is essential in any event to resolve 
the target from the signal from the ground or sea 
directly beneath the missile. This signal is always 
present, since the antenna system cannot be made 


completely directive. Indeed no radar-homing meth¬ 
od appears to be currently available if the missile-to- 
target distance is less than the altitude of the mis¬ 
sile, as it might be in some antiaircraft versions. 

Tracking both in azimuth and in range is accom¬ 
plished in the missiles of the Washington Project by 
so controlling the missile that it heads continuously 
for the target, i.e., on a pursuit course. A method 
which keeps the radar dish continuously pointed at 
the target and controls the missile by the departure 
of the axis of scan from some reference axis in the 
missile has been suggested and should be explored. 
This technique offers better possibilities for attacks 
along an interception course than does the method 
used thus far. 

To provide for failure of the signal due to the fluc¬ 
tuating character of radar or for other reasons, the 
servomechanism 15 must contain a memory element 
which will keep the missile on course until the signal 
is restored. Even with such an element in the servo¬ 
mechanism, the cone of vision of the radar receiver 
must be sufficiently broad to ensure that the target 
will be contained within it when the signal is restored. 

123 Aerodynamic Problems 

A missile has six degrees of freedom: translation 
along and rotation about the axes of roll, yaw, and 
pitch. Rarely will it reach a state of complete equi¬ 
librium. Bomb-like missiles are, in general, still under 
acceleration along the roll axis at the instant of im¬ 
pact. Lanchester * 1 2 3 has shown that a glider with con¬ 
trol surfaces fixed in the neutral position does not fly 
a straight line under the propulsion of gravity; 
rather, there is a periodic increase in speed with a 
corresponding increase in lift which flattens the glide 
path. The component of gravity in the direction of 
flight being reduced, the glider slows down, loses lift, 

b In this volume terms involving the word “servo” have a 
special significance beyond that usually ascribed to them in 
aircraft-engineering practice: 

1. Servo link , a mechanical power amplifier by which signals 
at a low power level are made to operate control surfaces 
requiring relatively large power inputs (e.g., a relay and 
motor-driven actuator). 

2. Servo system , a closed feedback loop comprising the in¬ 
telligence device, automatic pilot if any, the amplifying link, 
and the missile itself. 

3. Servomechanism , the feedback loop exclusive of the mis¬ 
sile itself. 

This convention follows the principle suggested by Warren 
Weaver. (See Fundamental Theory of Servomechanisms , LeRoy 
A. MacCall, with foreword by Warren Weaver, D. Van 
Nostrand Co., New York, 1945.) 




10 


GLIDE BOMBS 


and consequently assumes a steeper glide path until 
sufficient speed has been acquired for the cycle to 
repeat. Thus the actual glide path is a sinuous curve 
about a downward-inclined axis. Since the lift varies 
with the square of the speed and nonlinearly with 
the attitude of the control surfaces with respect to 
the wind stream—angle of attack—the curve is not 
a sine wave. Lanchester called it a phugoid. The 
phugoid period in seconds is about one quarter of the 
missile velocity in feet per second. Glide bombs have 
a phugoid period of about 88 seconds. Time constants 
of angular motion are considerably shorter, ranging 
from a fraction of a second for rotation about the roll 
axis to a few seconds for yaw and pitch. Efforts to 
anatyze any missile statically and to design a servo¬ 
mechanism based on the assumption of equilibrium 
conditions must fail. 

In the steady state, i.e., after equilibrium is estab¬ 
lished, a glide bomb with properly actuated control 
surfaces flies rectilinearly along a path which is in¬ 
clined downward by an angle whose tangent is the 
ratio of drag to lift and at a constant speed propor¬ 
tional to the square root of the wing loading. The 
glide bombs designed under this project have a 
maximum lift-drag ratio of approximately 7 and a 
flying speed in the vicinity of 230 mph. 

In conventional aircraft, change in lift is accom¬ 
plished by deflecting an elevator so that the wing 
and consequently the entire structure undergoes a 
change in angle of attack. Were such a design applied 
to a homing glide bomb, the axis of scan would be 
tangent to the life of flight for a single elevator set¬ 
ting only. In glide bombs using this type of control, 
attempts have been made to compensate for varying 
angles of attack by a feedback link from the elevator 
which adjusts the angle between the axis of scan and 
the axis of roll. Such compensation can be made ap¬ 
proximately valid under conditions of equilibrium, 
but are applied only with great difficulty under tran¬ 
sient conditions. As has been pointed out, a steady- 
state, static equilibrium rarely is obtained with mis¬ 
siles. 

In the glide bombs of the Washington Project the 
problem was solved by the use of elevons —full-span 
control flaps on the trailing edge of the wings. De¬ 
flection of the elevons varies the lift of the wing from 
approximately zero to a large positive value. Proper 
location of the center of gravity of the missile and a 
fixed tail structure provide moments from the down- 
wash which balance the moments produced by the 
flaps. The result is a missile which, in the steady 


state, maintains its angle of attack constant within 
1 degree for the normal operating range of the con¬ 
trols and within 4 degrees for full elevon excursion. 

In lieu of a rudder, differential operation of the 
elevons sets up a roll which in turn produces a com¬ 
ponent of wing lift normal to the line of flight. This 
lift, applied to the mass of the missile, produces a 
centripetal acceleration. 

The dynamic analyses of the performance result¬ 
ing from these methods of control are outlined in 
Sections 1.4 and 1.5. 

124 Interdependence of Controls and 

Correlation with Intelligence Coordinates 

The controls which affect the moments about the 
three axes of a missile are not independent. Thus 
rudder action produces a slight rolling moment in 
addition to the yawing moment which changes the 
heading of the missile to the right or to the left. 
There follows a larger rolling moment because of the 
reduced lift of the wing, whose absolute speed is de¬ 
creased by rotation about the yaw axis. Similarly the 
ailerons produce a yawing moment in addition to 
their primary roll moment. The roll moment result¬ 
ing from the yaw may either add or subtract from 
the roll produced by the ailerons directly. The inter¬ 
action of controls in pitch with yaw and/or roll are 
small. 

Homing devices and television are essentially two- 
dimensional intelligences, i.e., they give indication as 
to the azimuth and elevation of the target with re¬ 
spect to the axis of scan. Radar offers the possible 
addition of a third coordinate, the slant range, but 
no embodiment of this principle has been applied to 
missiles. The servo system has to take into account 
this discrepancy between two-dimensional control 
and the three-axis control problem which the preced¬ 
ing paragraph showed to be inherent in aeromissiles. 

Furthermore, roll of the missile, since the axis of 
roll in general is neither coincident with nor exactly 
parallel to the axis of scan, has an effect on the ap¬ 
parent position of the target with respect to the axis 
of scan and therefore on the error signal. Pitching 
and yawing have similar but reduced effects. These 
effects are subject to quantitative analysis in the 
idealized case but are difficult to generalize. (See 
Chapter 7 for a more complete three-axis develop¬ 
ment of the dynamics of the Pelican family of mis¬ 
siles.) 

The fuller discussion of these interactions between 



GENERAL 


11 


the elements of the servo system in Chapter 12 indi¬ 
cates the possibility of their reduction. Homing mis¬ 
siles can be made to work with the interaction pres¬ 
ent. Further aerodynamic and servo-system research 
looking toward the possible elimination of interaction 
is urged. 

1 - 2,5 Stability and Hunting 

Like any feedback loop, the servo system consist¬ 
ing of the homing missile and its controls will be un¬ 
stable if phase and gain relationships around the loop 
are not kept within appropriate limits. In general, 
instability will result in angular hunting about the 
three axes of the missile; in extreme cases of instabil¬ 
ity the missile may turn completely over and fall in 
a spin. A hit by a missile flying with a sustained hunt 
is purely fortuitous, depending upon the amplitude 
of the oscillation and its phase at the instant of im¬ 
pact; if the amplitude of hunt can be kept small, the 
miss will be small. 

The source of instability lies, as has been indicated, 
in the phase-gain characteristic of the loop. As the 
error signal crosses zero, the control surfaces should 
cross their neutral position. The lag between these 
events permits an input of energy into the system 
which must be absorbed by aerodynamic damping. 
The amplitude of hunt will build up until the cyclic 
energy input is equal to the energy dissipated in 
damping. 

Two methods of reducing the hunt have been ap¬ 
plied. An electric phase advancer can be designed 
which will reverse the error signal before the error is 
actually zero. Such an advance permits the control 
surfaces to pass their neutral position as the target 
comes on course, or even to lead that event. The 
latter situation corresponds to overdamping and is 
inherently nonoscillatory. This correction is secured 
at the cost of considerable attenuation of the error 
signal and requires a correspondingly increased gain 
in the receiver amplifier. A second method has been 
used with success. A rate gyro measures the angular 
velocity with which the missile comes on course. The 
indication of this gyro biases the flight-control equip¬ 
ment responsive to the radar so that the moment 
exerted on the missile toward true course depends not 
only upon the error in heading but also upon its first 
time derivative. By making the component of mo¬ 
ment which is dependent upon the derivative of error 
sufficiently large, the system is rendered nonoscil¬ 
latory. 


No aspect of the design of homing missiles is more 
important in achieving final accuracy than the prop¬ 
erty of dynamic stability. Serious study of Chapter 
12 and the references cited there is urged. 

12 6 Moving Targets and Wind 

If the target is in motion with respect to the mass 
of air through which the missile flies, the missile will 
attack the target along a pursuit curve. W. B. Klem¬ 
perer in Appendix A has shown that a missile with 
finite maneuverability cannot in general hit a point 
target unless the ratio of the speed of the missile to 
the speed of the target lies between 1 and 2. It is 
obvious that except for a level, head-on attack the 
missile will be unable to overtake the target if the 
ratio is less than unity. If the ratio is in excess of 2, 
except for a level, head-on attack or a stern chase, 
the curvature of the pursuit curve becomes infinite, 
requiring an infinite maneuverability on the part of 
the missile. 

Three-dimensional analysis of the pursuit problem 
is extremely difficult; no complete study of it seems 
to have been made. An idealized solution has been 
made for glide bombs under the direction of the 
Applied Mathematics Panel of the National De¬ 
fense Research Committee. It shows that against 
ship targets 2 the miss due to pursuit-course curvature 
is very small if the azimuth of the target at launching 
is about 145 degrees with respect to its path. 

A solution to the problem may lie in a computer 
which will cause the missile to lead the target and to 
fly a rectilinear interception course. A servo system 
within a servo system is implied, and a nice adjust¬ 
ment of the phase-gain relationship is required. Dry- 
den indicates in Chapter 12 that such a computer 
amounts to positive feedback and that instability 
may result. Doubtless, inept choice of parameters in 
the design of such a computer can lead to failure. 
The techniques learned in the development of fire- 
control equipment, however, provide powerful tools 
for the attack on this problem. 0 

Errors of alignment of the axis of scan with the 
tangent to the flight path give rise to errors analogous 
to those due to the pursuit curve. If the angularity 
between the line of flight of the missile and the axis 
of scan—the bore-sight error —is constant, the pro¬ 
jection of the trajectory is a logarithmic spiral with 

c A study of Aiming Controls in Aerial Ordnance by George 
A. Philbrick, STR Division 7, Volume 3, Part I, is recom¬ 
mended. 






12 


GLIDE BOMBS 


infinite curvature at the point of impact. With finite 
maneuverability in the missile, a miss proportional 
to its minimum turning radius and to the square of 
the bore-sight error will occur. So far as the Division 
is aware, no analysis of the combination of these ef¬ 
fects has been published; such an analysis is seriously 
needed. In this project the effect has been minimized 
by reducing as far as possible variations in angle of 
attack and by careful alignment of the axis of scan 
with a fixed reference axis in the missile to avoid 
bore-sight error. 

!■* AERODYNAMIC FEATURES 3 

The glide bombs developed under this project con¬ 
sist of high-wing monoplanes of laminated-wood, 
monocoque construction. The wing loading is ap¬ 
proximately 70 psf, and the three sizes (8 ft, 350 lb; 
10 ft, 1,000 lb; and 12 ft, 2,000 lb) are substantially 
homologous. (See Figure 1.) 

Control is obtained from elevons which operate in 
conjunction for control in pitch and differentially for 
control in roll. Turning moments are derived from 
roll. A fixed empennage with twin disk-shaped fins 
stabilizes the missile in yaw and pitch; the position 
of the empennage with regard to the center of gravity 
of the whole assembly and the decalage between the 
wing and stabilizer provide nearly constant angle of 
attack. 

The model with the 8-ft wing span was tested at 
an airspeed of 90 mph in the NACA wind tunnel at 
Langley Field. The following tabulation presents the 
significant characteristics of the structure at trim. 
The data were obtained by allowing the missile to 
assume an attitude in the plane of the yaw and roll 
axes such that the pitching moment (C m ) was zero. 
Direct observations of change in attitude gave the 
variation in a, the angle of attack. Lift and drag were 
measured in the usual manner. Longitudinal stability 
at trim (dC m /da when C m = 0) is not directly re¬ 
ported as a function of elevon angle, 8. It is deter¬ 
minable, however, from the data reported, and Sec¬ 
tion 1.4 shows the important application of this func¬ 
tion to the study of longitudinal hunting. 

The flattest glide angle (8.0 degrees) which is de¬ 
rived from a lift-drag ratio of 7.11 is equivalent to 
1.35 miles of range for each 1,000 ft of descent. Ex¬ 
tension of the operation to the absolute limits leads 
to failure, however, since the departure from linearity 
of the Cl/Cd vs 8 characteristic is great, and more 
particularly because the angle of attack is not con- 


Flap 

angle 

8 

(degrees) 

Lift 

Coefficient 

Cl 

Drag 

Coefficient 

C D 

Steady-state 
glide angle* 
cot -1 Cl/Cd 
(degrees) 

Angle of 
attack 

a 

(degrees) 

-20 

0 

0.062 

90 

4.1 

-15 

0.03 

0.053 

60.4 

3.7 

-10 

0.08 

0.048 

30.9 

3.5 

- 5 

0.15 

0.046 

17.1 

3.5 

0 

0.23 

0.048 

11.8 

3.8 

5 

0.34 

0.054 

8.9 

4.4 

10 

0.46 

0.066 

8.1 

5.3 

15 

0.59 

0.083 

8.0 

6.2 

20 

0.72 

0.110 

8.8 

7.5 


*Referred to horizontal flight. 


stant. Within the range 8 = — 10 degrees to 8 = +5 
degrees, the variation in a is only 0.9 degree. In 
practice the missile was dropped at a range to require 
a glide-path ratio of 3.5 to 4.0 (5 = — 4 degrees to 
8 = +5 degrees). In the region reasonably adjacent, 
Cl/Cd vs 8 is essentially linear. 

While the full excursion of the elevons to the posi¬ 
tion of flattest glide (8 = 15 degrees) cannot be suc¬ 
cessfully utilized for stretching the operating range 
of the missile, the implication that mechanical limits 
can be set on their movement (say at —10 and +7 
degrees, which represent the limits of approximate 
linearity of L/D vs 8) would be wholly false. Since a 
differential elevon action is superimposed on joint 
elevon action for combined turn and change of glide- 
path ratio, the full excursion may be required. 

No tests were made to determine a with C m = 0 
when combined differential and joint displacement 
of the elevons is applied. It may be that there is 
concealed here a considerable variation in angle of 
attack which vitiates the accuracy of the missile. 
Further studies appear to be indicated to explore 
this field. Moreover the dynamic application of 
steady-state data from the wind tunnel is a matter 
which strongly demands further inquiry. 

1.4 longitudinal stability 
OF GLIDE BOMBS 

141 Analytical Studies 4 

The stability characteristics of a glider are those 
qualities which determine its motion after a small 
deviation from an initial condition of equilibrium. 
Longitudinal stability limits these properties to those 
which determine translation along the axes of roll 
and yaw and rotation about the axis of pitch. It is 
the group of characteristics which, when taken to- 







LONGITUDINAL STABILITY OF GLIDE BOMBS 


13 


gether with those of the homing device, the automatic 
pilot, and the servomechanism, defines the perform¬ 
ance of a homing glide bomb in the range sense. 

The conditions of equilibrium in flight are defined 
by: 

W sin 7 + hf>SV 2 C D = 0 
W cos y + \ P SV*C L = 0 (1) 

\pSV 2 C m = 0 

where 5 W is the weight of the missile, 

S is the effective wing area, 

V is the velocity of the missile, 
p is the density of air, 

7 is the downward inclination of the flight 
path, and also 
m is the mass of the missile, 

B is its moment of inertia in pitch, 
q is the velocity head, 

Q is the attitude of the missile, and 
c is the mean aerodynamic chord. 

AfterVsmall displacement in V, 7 , Cd, Cl, and C m 
—as by operation upon them by the increment A— 
these equations become: 

W cos 7 A 7 -|- § pSV 2 ACd H” pSVCdAV = mV 
- W sin 7 A 7 + ipSV 2 AC L ~ pSVC l AV = mV 7 (2) 
pScVCmAV + i P SV 2 cAC m = Bd 

In his classical solution of these equations Lan- 
chester 1 considered the glider as flying in a conserva¬ 
tive field, i.e., no drag, and considered only the case 
where the control surfaces are fixed. In the present 
study the development is extended to include the 
major effects of drag and the motion of the elevons 
which comprise the control surfaces of the glide 
bombs of the Washington Project. In arriving at an 
analytical solution which defines completely the mo¬ 
tion of the missile in the plane of the yaw and roll 
axes the following simplifying assumptions are made: 

1. Angular velocities and accelerations have a neg¬ 
ligible effect on Cl and Cd, and therefore 

2 .Cl and Cd are functions of a and 8 only. 

3 . The relationship between Cl and Cd on the one 
hand, and a or 8 on the other is linear. 

4. Aa, AV, q, a, 8, and 8 are considered small but 
appreciable. Terms containing products of two or 
more of them are considered negligible. 

5. Static characteristics as measured in the wind 
tunnel are considered to hold in the dynamic or 
transient state. 

6 . The servo system is ideal, i.e., linear and with¬ 
out phase distortion. 


Now if: 

AV 
V V 

L = acceleration due to lift = —— - L 

m 

D = acceleration due to drag = - 2 


M = angular acceleration in pitch = - 
L a = 

M a = 

Ma = 


B 


6L 

T 

_ dL 

da ’ 


68 

6M 

M 

dM 

da ’ 

Ivld 

68 

dM 

Mi 

dM 

da ’ 

68 


dq 


we obtain: 


(3) 


LAy + 2 Dv -T D a Aa -f- Ds8 -|- Vl — 0 
— DAy T - 2 LV T - L a Aa T - Ls8 — V’ = 0 
M q { 7 + a) + M a AoL + M&8 + M^a + M's8 

— 7 — a = 0 


(4) 


There is nothing in equations (4) that requires 
linearity or absence of phase distortion in the servo 
loop. Such limitation is required for the analytical 
solution. The limitation can be expressed: 

- 8 = K(A0 + c'0) (5) 

where K and c' are factors of proportionality. Equa¬ 
tion ( 5 ) is a fundamental equation of an idealized 
servomechanism; it requires that the elevon dis¬ 
placement be proportional to the error in heading 
and to its first time derivative ." 1 After inserting equa¬ 
tion (5) in equation (4) a solution is obtained which 
gives, independently of the value of K, a short- 
period, highly damped oscillation. This is the so- 
called rapid incidence adjustment , and its independ¬ 
ence of K shows it to be unaffected by servo design. 

In addition to the short-period oscillation, there is 
a motion whose character is dependent upon K. If 
K is small, this motion is a long-period oscillation 
similar to the phugoid of Lanchester but damped; 
if K is large, the motion is not oscillatory and may 
be either regenerative or degenerative. 

If (d)(L/D)/(d8) > 0, then the motion will be de¬ 
generative or stable; if ( d ) ( L/D)/(68) < 0 , instability 
may result. Such a situation sometimes arises when 
the glide bomb is launched at too low a speed or 

d See Chapter 4 for a further discussion of the idealized 
control regime for a missile guided by full-span flaps. 







14 


GLIDE BOMBS 


when too flat a glide angle—excessive range—is at¬ 
tempted. 

142 Solution by an Electromechanical 
Analogue 

The establishment of the equations of motion of a 
glide bomb and their solution for the case where 
control is effected through an ideal servomechanism 
led to a broader study with more realistic servo¬ 
mechanisms through the use of an electromechanical 
analogue. This technique proved so powerful in at¬ 
tacking such problems that it is discussed at some 
length in Chapter 11. The applications of the prin¬ 
ciples of dynamic similitude to the design and the 
details of construction of the table are discussed 
there. 

The analogue consisted of a table free to rotate 
about a vertical axis but with damping provided for 
that motion (see Figure 2). The inertias of the ro¬ 
tating system, the damping factors, and the com¬ 
pliances are adjusted so that oscillation about the 
vertical axis corresponds to the pitch oscillation of 
the missile about its horizontal axis. A pitch gyro 
(see Section 1.6) is mounted on the table, and its 
output fed to a computing circuit. The computing 
element in turn actuates the servo link, whose out¬ 
put is spring-loaded to correspond to the hinge mo¬ 
ment of the elevons. 

The output of the servo link, both as to velocity 
and displacement, is returned via the computing cir¬ 


cuit as a feedback to the table. Thus the system com¬ 
prises a closed loop with all the elements of the mis¬ 
sile system except the automatic homing device. 
Several types of gyros and strengths of bias coils (see 
Section 1.6) were tested, together with a very large 
number of servo links. Typical curves are shown in 
Figure 3. 

1 5 LATERAL STABILITY 

OF GLIDE BOMBS 

The problem of lateral stability is not so simple as 
that of longitudinal stability. The reason is suggested 
in Section 1.2.4, where the interrelation between 
aerodynamic forces and moments was described. 
Three principal modes of motion are considered: side¬ 
slip velocity, roll velocity, and yaw velocity. These 
are closely interrelated and are in turn sensitive to 
elevon displacement, attitude in pitch, and inclina¬ 
tion of the flight path. 

A complete analysis of the problem is beyond the 
scope of formal mathematics although a study with 
simplifying assumptions can be made rewarding; in 
Appendix B Skramstad has made such a study. Even 
with the assumption of linearity of response, constant 
air density, and the overwhelming preponderance of 
certain design properties such as roll damping, the 
dynamic equations lead to a quintic for the evalua¬ 
tion of the roots. 

In two-winged missiles the roll damping is large. 
Unless the servo system introduces a positive feed- 




DAMPED TURNING TABLE 


\ 


SERVO LINK UNDER TEST 


GYRO UNDER TEST 


Figure 2. Flight-test table and accessories. 





LATERAL STABILITY OF GLIDE BOMBS 


15 




40 50 60 70 80 

TIME IN SECONDS FROM RELEASE 


Figure 3. Longitudinal oscillation of Pelican in flight. 


back, the assumption of roll-damping predominance 
is valid. Under this circumstance the motion of the 
missile after a perturbation in roll, yaw, or sideslip 
breaks down into three components: 

1. A rapid exponential subsidence of the disturb¬ 
ance which, for the Pelican family of missiles, has a 
time constant of about 1 second. 

2. A slower exponential factor, which may be either 
regenerative or degenerative. This component ex¬ 
presses the spiral stability of the missile: if it is de¬ 
generative, the missile is spirally stable; if it is re¬ 
generative, the missile is spirally unstable. The glide 
bombs of the Washington Project were spirally 
stable, although the margin of such stability was 
small. 

3. An oscillation whose period is determined largely 
by a yawing moment due to sideslip and whose damp¬ 
ing is determined largely by a yawing moment due to 
rate of yaw. The Washington Project’s missiles have 
a period of oscillation in yaw of approximately 1 
second and a time constant of about 1.67 seconds. 

Simulation was invoked in the analysis of the air¬ 
frame-automatic-pilot links of the servo-system 
chain. In the early phases of the program an analogue 
model was constructed wherein roll and yaw were 
studied separately, with corrections applied to the 
results to take their interrelation into account. In the 
final work the entire servo system including the hom¬ 
ing intelligence was simulated. (See Chapter 7.) 


The necessity of an automatic pilot (see Section 
1.6) had been established. Its performance about the 
roll and yaw axes was dictated by a rate gyro whose 
axis of rotation was inclined to the roll axis of the 
missile so that it was sensitive to both rate of bank 
and r^te of turn. Thus roll and bank of the complete 
servo system were doubly interrelated: through the 
fundamental aerodynamic properties of a two-winged 



Figure 4. Radar homing signal versus error in 
heading. 





























































16 


GLIDE BOMBS 




V) 

ft+5 

oc 



1 

Bright 

A. A l 

\AAA 

\ A A A 

. . A A 

A A A / 

|A A . 



aA/V 


^vvv 

V^Aa. 

V 

Ileft 


/\r SJ 

V v V ) 

v Vv N 

A/Vv 

J \r 'rv 

^vv\a 

MAai 

Vwv 

v*Vr V 





















0 5 10 15 20 25 30 35 40 45 50 55 60 65 60 

TIME IN SECONDS FROM RELEASE 


Figure 5. Output curves of flight-test table compared with flight tests, roll and yaw. 


missile and through the coupling of the rate-of-bank- 
and-turn gyro. 

The coefficient of coupling effected by the gyro 
depends upon the inclination of its rotational axis to 
the roll axis. This angle of inclination was varied to 
determine its optimum value. Runs were made with 
the inclination angle at values intermediate between 
2.5 degrees and 20 degrees. At 2.5 degrees the yaw 
oscillation was negatively damped; at inclinations of 
the gyro greater than 20 degrees the system was over¬ 
damped and too sluggish. A value somewhat under 
15 degrees was found most satisfactory in flight tests. 

The radar-homing equipments (see Section 1.7) 
were designed to give an error signal which is propor¬ 
tional to the error in heading for relatively small 
errors and is constant at larger errors. The error angle 
at which the signal saturates is adjustable (Figure 4). 
Exploration of the effect of varying the saturation 
angle, called radar width, on the simulative flight 
table showed that a radar width of ± 6 degrees was 


most suitable (Figure 5). This value also gave the 
most satisfactory performance in flight. 

16 AUTOMATIC PILOTS 6 

1,6,1 General 

The initial flight tests on this project were made 
without automatic flight-control equipment. The ex¬ 
perience of toy-model builders was a strong force 
leading to the belief that automatic stabilization 
would not be required. This hope proved to be false 
in spite of the inherent stability expected from the 
high-wing design. 

The fundamental decision to be made in the design 
of the automatic pilot was the choice between free 
gyros and rate gyros. The free gyro provides a ref¬ 
erence in space which—if the frictional forces in the 
pivots and take-offs are minimized—remains well 
fixed. Thus the datum from which a free-gyro-con- 





















































































RADAR-HOMING CONTROL—RHB CONTROL 


17 


trolled automatic pilot controls is absolute, and the 
whole assembly possesses memory. Loss of control 
signal, either from the radio-control link or from 
temporary failures of an automatic homing system, 
will not result in departure from course beyond the 
tolerances of the gyro. 

In the rate gyro the problems of take-off friction 
are very easily solved since the whole structure is 
spring-restrained about the axis where movement re¬ 
sults in contact closure. It is relatively simple to 
make contact friction small compared with spring 
tension. As the free gyro holds the attitude of the 
missile fixed, so the rate gyro holds its first time de¬ 
rivative fixed. If the rate gyro could have a zero 
tolerance or dead band, an ideal situation, the results 
would be identical. In a practical embodiment, the 
rates of departure from desired attitude can be suc¬ 
cessfully small if the control signals are reasonably 
continuous. 

These considerations, together with the consider¬ 
ably more plentiful supply of rate gyros, led to their 
selection. It is interesting (see Chapter 7) that a suc¬ 
cessor investigator reversed this decision, free gyros 
having in the meantime become much more readily 
available. 

1-6,2 Basis of Design 

From the outset two rate gyros were used. One has 
its rotor axis parallel to the axis of yaw and is free 
to precess, against spring restraint, in response to 
pitching; its action is thus substantially that of a 
rate-of-climb indicator. The other gyro is mounted in 
the plane of the yaw and roll axes, with its axis in¬ 
clined to the latter axis by approximately 15 degrees 
(Figure 6); it is thus primarily sensitive to rate of 
turn but also has a secondary sensitivity to rate of 
roll. Quantitatively the effectiveness of the instru¬ 
ment as a rate-of-turn indicator is 97 per cent and a 
a rate-of-bank indicator 25 per cent. 

Electromagnets actuate the gimbals to close con¬ 
tacts which control the servo link. Thus if the “down” 
contacts are closed by its magnet, the elevons are 
deflected upward (the direction of decreasing lift) 
until the rate of pitch produces a precessional torque 
on the gimbal sufficient to overcome the magnetic 
pull. 

Similarly if the “left” contacts are closed by sole¬ 
noid action, the left elevon is deflected upward and 
its mate downward until the rate of roll and the rate 
of turn, biased respectively by the sine and cosine 


of the angle of inclination of the gyro rotor, produce 
a precessional torque sufficient to overcome the mag¬ 
net force of the gyro coils. 

On loss of control signal in pitch, the missile will 
be restrained, except for gyro tolerances, to zero rate 
of pitch and will therefore continue along a substan¬ 
tially constant glide-path angle. On loss of control 
signal in azimuth, the turn-sensitive component of 
precessional torque will right the missile at a rate 
permitted by the roll-sensitive component. There¬ 
after the glider will—again neglecting tolerances in 
the gyro—fly a straight course. 

In the television version of the glide bomb, relays 
in the output circuit of the radio-control receiver 
energized, at a fixed current value, appropriate elec¬ 
tromagnets for “up,” “down,” “right,” and “left” 
control. Thus the bombardier had the option of im¬ 
parting to the missile fixed rates of correction in range 
and in azimuth. The amount of correction was gov¬ 
erned by proportioning the ratio of time-on to time- 
off. Skilled pilots did not always make the most suc¬ 
cessful control bombardiers. 

The output of the homing radar receiver is single¬ 
valued in current as a function of error through a 
considerable range. At a “left” error of about 20 
degrees, the output is 8.0 ma direct current. With a 
20-degree “right” error, the output is equal in the 
reverse sense. For approximately 6 degrees adjacent 
to zero, the response is linear from —8.0 ma to +8.0 
ma. Thus, under radar control the servo system 
operates to give a rate of return to true course pro¬ 
portional to the error in heading. In order to reduce 
yaw hunting, a third gyro sensitive to rate of yaw 
was later added to make the system inherently non- 
oscillatory. All three gyros are shown in Figure 7. 

RADAR-HOMING 
CONTROL—RHB CONTROL 7 - 8 

1-7-1 General Principles 

The all-weather property of a microwave makes it 
an attractive agency for the control of a guided mis¬ 
sile. The directional property of radar (Radio Direc¬ 
tion and Range) suggested, even before the formation 
of the Division, that the signals reflected from a 
water-borne target could be used to guide a pilotless 
aircraft to impact with exact accuracy. The radar¬ 
homing bomb [RHB] was conceived as applicable to 
powered drones, gliders, and to glide bombs. Within 



18 


GLIDE BOMBS 


Y 



the Division, only its application to glide bombs of 
the Pelican family was prosecuted. 

Military targets in general do not emit radio waves; 
it is, therefore, necessary to illuminate them with ra¬ 
dio radiation if a missile is to home on them through 


the agency of radar. In RHB the target is illuminated 
by the standard radar search equipment in the car¬ 
rying aircraft. The equipment consists of a powerful 
transmitter which emits a narrow beam of short¬ 
wave trains of very high intensity. The reflections 



Figure 7. Block diagram of glide-bomb control system. 




































































RADAR-HOMING CONTROL—RHB CONTROL 


19 



from the target are picked up by a receiver in the 
missile. The output of the receiver actuates the auto¬ 
matic pilot (Figure 8). The wave trains or pulses 
occur at a frequency of about 800 per second and are 
0.7 microsecond long. The illuminating radiation 
thus consists of a series of pulses in space, each 740 
ft long and separated by 233 miles. The quasi-optical 
property of the radar frequency, 3,000 me, provides 
directional information as to the target’s location, 
and the pulsed character of the illumination provides 
means of determining its range. 

The energy radiated from the illuminating trans¬ 
mitter is focused into a narrow beam (approximately 
+ M degree to half-intensity) by the antenna struc¬ 
ture. Within the field thus illuminated, every object 
returns some reflected or scattered energy; the study 
of the energy reflected from various types of targets 
and from their typical backgrounds is basic to the 
design of successful radar-homing systems. 

17 2 Target Contrast 

Although all military targets will return a radar 
echo, it may be impossible for the receiver to resolve 
the desired signal from the multiplicity of echoes 
from the objects which surround the target. This is 
especially true of land targets. Even for water-borne 
targets there may be considerable clutter from the 
ocean near the target, particularly when an appre¬ 
ciable sea is running. This is in addition to the very 
large “altitude” signal which is returned from the 


ground or sea directly beneath the receiver and which 
the receiver must be designed to reject. 

In the microwave range the resolving power of a 
radar receiver increases with increasing frequency. 
Qualitative studies were made by the Radiation Lab¬ 
oratory. The findings in the 3,000-mc band indicated 
that the resolving power of the receiver was increased 
if the incidence angle of the illuminating beam was 
high. Extension of these studies by the MIT group 
under Division 5 was made against a ship illuminated 
at depression angles (the complement of incidence 
angle) between 2 and 11 degrees. These tests showed 
that nearly all the energy scattered from the ship 
target lies in the sector extending from the sea to a 
plane some 45 degrees above the incident beam. The 
energy is distributed in azimuth on the illuminated 
side of the ship in a roughly uniform pattern. Homing 
on a reflected signal at approach angles steeper than 
50 degrees is therefore marginal; above 70 degrees it 
is impossible. 

These studies were extended and made more quan¬ 
titative in the 10,000-mc band. The energy returned 
from the target and from the clutter was evaluated 
in terms of the cross-sectional area of a spherical re¬ 
flector which would return the same echo. This area 
is defined as 

_ 16tt *nWPr 

~ P t G t A r 

In the foregoing: 

ri is the range from the receiver to the target. 

r 2 is the range from the transmitter to the target. 




20 


GLIDE BOMBS 


P r is the peak received power. 

P t is the peak transmitted power. 

G t is the transmitter antenna gain. 

A r is the effective area of the receiver antenna. 

Values of S a were determined for targets ranging 
in size from LST’s to 10,000-ton Liberty ships and 
for the sea clutter through a broad range of rough¬ 
ness. The receiver and transmitter in these experi¬ 
ments were at the same location. As a function of 
depression angle the equivalent cross-sectional areas 
are: 


Depression 
Angle (degrees) 
20 
30 
40 
50 
55 
60 
70 


Equivalent Cross-Sectional Area (sq ft) 
Target Sea Clutter 


10 4 to 10 5 
10 4 to 10 5 
10 4 to 6.5 X 10 4 
10 3 to 2 X 10 4 
250 to 10 4 
Negligible 
Negligible 


Negligible 
Negligible 
7 to 1.5 X 10 3 
15 to 1.6 X 10 3 
40 to 5 X 10 3 
200 to 1.6 X 10 4 
4 X 10 3 to 6.4 X 10 4 


Similar studies are required for land targets. 

Successful homing operation is not to be expected 
at depression angles close to 50 degrees, the limit of 
resolution from clutter given in the foregoing table. 
The fading character of radar will make range track¬ 
ing at such angles of dubious reliability. 


1,7,3 RHB Receiver 

The energy returned from the target is detected 
and measured by a superheterodyne receiver located 
in the missile. The local oscillator consists of a kly- 



Figure 9. Block diagram of RHB receiver. 


stron tube with a tunable cavity. The signals from 
the antenna and from the local oscillator are com¬ 
bined in a crystal mixer to produce heterodyne beats 
which are then amplified. The seven-stage i-f ampli¬ 


fier operates at 30 me; it is tuned to this frequency 
but with sufficient band-pass to prevent undue dis¬ 
tortion of the 0.7-microsecond pulses. A detector 
passes the envelope of the pulses which is amplified 
and, after operation of the tracking system, com¬ 
mutated to provide directional data to the automatic 
pilot (Figure 9). 

Fundamental information as to the direction of the 
target is obtained by a scanning antenna. A disk 
dipole is located at the focus of a 12-in. diameter 
parabolic reflector of 3.6-in. focal length. The “beam- 


90 * 



SCANNED FIELD OF VIEW 



RECEIVED SIGNAL PULSES 

Figure 10. Conical scanning process. 


receptor” character of this antenna system yields a 
signal which is down 50 per cent when the source is 
+ 11 degrees from the axis of the reflector. The re¬ 
flector rotates at 1,800 rpm about an axis which is 
inclined to its optical axis by 5.5 degrees. Thus, dur¬ 
ing one revolution of the antenna 26% pulses are re¬ 
ceived from the target. If fading is neglected, the 
signal strength during a scanning cycle will depend 
upon the bearing of the target with respect to the 
axis of rotation of the reflector—the axis of scan. 
For a target on the axis of scan the returned signal is 
constant; for small error angles the variation in signal 
strength over a scanning cycle is approximately 
sinusoidal. The envelope (Figure 10) is integrated 
over quarter-cycles, and the total received energy in 
diametrically opposite quadrants, when amplified, 
operates to bias the rate gyros as shown in Section 
1 . 6 . 













































































RADAR-HOMING CONTROL— RHB CONTROL 


21 


Within the illuminated field there will generally be 
several reflectors. For a horning attack to be success¬ 
ful one target must be selected and the attack pushed 
home on it. One method of eliminating all but the 
desired signal would be to make the beam sensitivity 
of the receptor extremely narrow. This would require 
an antenna reflector too large to be carried in the mis¬ 
sile and might result in complete loss of the target 
when the motion of the missile is made erratic by 
gusts. Pulsed radar provides a practicable method 
for range tracking. 

Range discrimination is essential in any event in 
radar-homing glide bombs because of the large 
amount of energy returned from the ground or sea, 
which is always the biggest target in the field. At 
ranges of 5 miles or more, and with the smaller an¬ 
tennas required for missiles, as much energy is re¬ 
ceived from the ground or sea directly below the mis¬ 
sile as from an average target; this altitude signal 
must be rejected by the receiver. In glide bombs, 
since the altitude is always less than the range to the 
target, a range discriminator makes it possible to 
eliminate the altitude signal. The target selector pro¬ 
vides range discrimination by “gating” the desired 
signal, i.e., allowing the differential amplifier to re- 



Figure 11. Block diagram of range-tracking circuit. 


ceive signals only at the exact time that echoes from 
the desired target are received. It also, through a 
memory circuit, adjusts the time of operation of the 
gate circuit to the range of the target, i.e., to track 
in range. 

Synchronization is accomplished by starting the 
tracking circuit (Figure 11) with the reception of the 
main pulse from the transmitter. The receiver nor¬ 
mally is at full sensitivity. Upon reception of the 
main pulse via the rear-view antenna a negative 
pulse is applied to the grids of the first and second 
stages of the i-f amplifier; this biases the receiver 


below the noise level and prevents the reception of 
any signal until the first echo. This main gating pulse 
is about 200 microseconds wide. 

Within the main gate are two narrow gates, each 
0.6-microsecond wide with a 0.2-microsecond overlap 
yielding a total gate of 1.0 microsecond. The double 
gate is tripped by a single-cycle multivibrator at a time 
after the main pulse equal to the transit time from 
the missile to the target and back. The time of trip¬ 
ping is determined by the rate of discharge of a con¬ 
denser in the memory circuit. Were the range to de¬ 
crease at a constant rate, it would be necessary only 



Figure 12. Operation of double-tracking gate. (A) 
phaseable pulse tracking too slowly—excess of energy 
in Gate No. 1; (B) phaseable pulse tracking too fast— 
excess of energy in Gate No. 2; (C) phaseable pulse 
correctly synchronized—equal energy in both gates. 

to have the double gate traverse the wide gate at a 
constant rate. The function of the double gate is to 
adjust the tracking rate so that the target echo will 
always be centered within it (Figure 12). If the rate 
of tracking is too fast, the portion of the echo lying 
in the second 0.6-microsecond gate will exceed that 
in the first, and the voltage against which the multi¬ 
vibrator condenser discharges will rise, and vice versa. 
With 3,000 me there are 2.1 adjustments of the mem¬ 
ory circuit in each 0.7-microsecond pulse. Time con¬ 
stants of the memory circuit are such that, should 
fading cause loss of synchronization, the memory cir¬ 
cuit will track at the correct velocity for several 
seconds. 

A second function of the narrow gate is to apply 
an AGC voltage to the i-f amplifier. This voltage is 
so adjusted that a strong signal is held to half the 
saturation level of the output stages. The AGC must 
not be so fast that the 30-c modulation will be ob¬ 
scured with consequent loss of directional informa¬ 
tion. AGC over a range of about 5.3 db is required 
to accommodate the increase in received power due 
to the operation of the inverse square law as the mis¬ 
sile approaches the target. 
























































22 


GLIDE BOMBS 


174 Maximum Range 

RHB receivers (Figure 13) have a threshold sensi¬ 
tivity of about 10 -12 watt; 10 -10 watt gives a reason¬ 
able margin for satisfactory range tracking. The rela¬ 
tion between range and received power is given by: 

P t D t 2 D r 2 S a 
nr2 64P r \ 2 

where r\ is the range to the transmitter in feet; 
r 2 is the range to the receiver in feet; 

P t is the transmitter output in watts; 

P r is the echo signal strength in watts; 

D t is the diameter of the transmitter antenna 
reflector in feet; 

D r is the diameter of the receiver antenna re¬ 
flector in feet; 

S a is the equivalent spherical cross-sectional 
area of the target in square feet. 

X is the wavelength in feet. 


At the dropping point r i and r 2 are identical. With 
airborne search equipment the transmitter output is 
approximately 30 kw. D t and D r are 2.5 ft and 1 ft 
respectively. With 3,000-mc radar—X about }/& ft— 
a safe maximum range for RHB against a Liberty 
ship is about 27 miles. 

175 Tests with RHB 

The basic development of RHB was carried out by 
MIT Radiation Laboratory under Division 14. Lab¬ 
oratory prototypes were procured by them for initial 
tests in powered aircraft and in glide bombs (Peli¬ 
can). When procurement by the Navy Bureau of 
Ordnance was instituted, the project was turned over 
in its entirety to Division 5. A few of the personnel 
associated with the project formed the nucleus of a 
new MIT radar group operating at the National 
Bureau of Standards under Contract OEMsr-240. 
Too much credit can hardly be proffered to Division 



POWER SUPPLY 


BLOWER 


REAR ANTENNA 


JUNCTION BOX 


RECEI 


VER 


BATTERY 


Figure 13. Receiver and power supply for RHB Pelican. 













RADAR-HOMING CONTROL—SRB CONTROL 


23 


14’s group for boldly undertaking a difficult venture 
and completing the basic research in approximately 
19 months. 

Early tests on RHB were in an AT-11 airplane 
with a human servo link. The target area was il¬ 
luminated by a large truck-borne radar search equip¬ 
ment stationed on a hill overlooking a portion of 
Boston Harbor. The output of the RHB equipment 
lighted four signal lights by which the pilot flew the 
airplane at selected ship targets. To passengers the 
accuracy seemed great; cold analysis of the motion 
picture records indicated that the airplane was not on 
true course for a high percentage of the time. 

Neither the Division nor the Navy realized the 
extent of the work remaining to be done. The entire 
problem of adjusting the response rates of the air¬ 
frame, the automatic pilot, and the radar equipment 
remained to be solved. 

When laboratory model-shop prototypes became 
available, an experimental range was established by 
the National Bureau of Standards and the Navy 
Bureau of Ordnance. Here testing continued against 
beacon radiators and reflectors suspended from bar¬ 
rage balloons. Faulty construction of the model-shop 
equipment masked the results of the tests. Instru¬ 
mentation (see Chapter 8), though well conceived, 
was not always adequate to establish conclusively 
which links of the servo loop failed. 

It was not until a complete radar laboratory was 
established and staffed with some twelve scientists 
and engineers, supported by technicians made avail¬ 
able from the Navy, that a measure of consistent 
success was obtained. Defects in the radar equip¬ 
ment which seemed trivial in the laboratory were 
wholly frustrating in the field. A third gyro sensitive 
to rate of yaw had to be provided to damp out oscil¬ 
lations. Everyone associated with the project had to 
learn that the automatic control of a glide bomb is 
a more difficult task than the combined operations 
of interpreting a radar scope and flying an airplane. 

Ultimately, consistent scores of 50 per cent hits 
from 11 miles were attained against a 10,000-ton 
ship. 

17 6 Self-Synchronous RHB 

The original RHB used the signal received directly 
from the transmitter as the synchronizing impulse. 
In all RHB’s, precautions must be taken to insure 
that sufficient energy is received. When the target is 
illuminated from an aircraft, the missile may not re¬ 


main within the beam because of evasive action of 
the aircraft or because of the difference in the speeds 
of the missile and aircraft. Hence the transmitter 
must be equipped with an additional nondirectional 
(or at least a very-broad-beam) antenna to ensure 
the radiation of some energy in the direction of the 
missile. Likewise the directional characteristics of the 
receiver antenna are such that little energy is re¬ 
ceived from the rear, which is the direction of the 
transmitter. It is necessary, therefore, to provide a 
rear-view antenna on the missile to ensure that suf¬ 
ficient energy is received directly from the trans¬ 
mitter to synchronize the range selection gate. 

The second system derives its synchronization 
from the reflected signal; hence the equipment is 
termed self-synchronous. Actually a very stable low- 
frequency oscillator is built into the receiver and an 
identical oscillator is used to control the pulse rate 
of the transmitter. The frequency of the receiver 
oscillator is automatically adjusted to that of the re¬ 
ceived pulse rate. The self-synchronous RHB was 
developed primarily because it offered a means of 
securing homing intelligence all the way into the 
point of impact with the target. Both the Pelican and 
the Bat (see below) experience difficulties at close 
ranges because of the coincidence of the initial pulse 
with the returning echo. Pelican requires the recep¬ 
tion of the direct transmitted pulse for synchroniza¬ 
tion, and Bat cannot make use of the echo when its 
transmitter is operating. The self-synchronous RHB, 
however, derives its synchronization from the echo 
only, and it can exclude the initial pulse, which is 
transmitted from many miles behind the receiver. A 
few experimental models of radar-homing equipment 
with this method of synchronization have been built 
and tested. 

18 RADAR-HOMING CONTROL— 

SRB CONTROL 9 

The receiving system used in RHB was used sub¬ 
stantially unchanged in the send-receive bomb [SRB] 
(Figure 14). The basic change was that the missile 
Bat carried the transmitter as well as the receiver 
(Figure 15). The whole equipment was developed by 
the Bell Telephone Laboratories under direct contract 
with the Navy Bureau of Ordnance. The Division 
rendered consulting service through its radar group 
established for the engineering development of RHB. 

Successful operation was required at ranges of 7 
miles with continuing tracking in to 100 yd. The 



24 


GLIDE BOMBS 




fourth-power law that such a system follows requires 
an AGC system with a range of 120 db against large 
cargo vessels; this requirement was met only with 
difficulty. 

AGC was applied from the second and following- 
stages of the i-f amplifier, the first stage being omitted 


to avoid pickup of noise on the AGC leads to the 
grid. The high loads on the first stage at close ranges 
gave rise to shock-excited spurious signals which per¬ 
sisted as “ringing,” distorting the true signal in the 
tracking gate. The frequency of the transmitter 
drifted with altitude changes, and the directional 



NOSE COVERING 


TRANSMITTER RECEIVER 


POWER .SUPPLY 


INVERTER 


Figure 15. SRB—Bat radar equipment. 





SERVO LINKS 


25 


response of the equipment was distorted because of 
change in loading impedance as the reflector rotated. 

The ghost signals arising in the i-f amplifier were 
eliminated by an engineering compromise in the Q 
of the circuit so that the ringing would be more 
damped, still leaving sufficient band-pass to avoid 
distortion of the 0.7-microsecond pulse. The master 
oscillator was pressurized to prevent frequency drift 
with altitude changes, and equipment within the mis¬ 
sile was rearranged to eliminate the squint caused by 
change in output impedance during scanning cycle. 

As in the case with RHB, defects in the equipment 
which seemed of little importance in the laboratory 
loomed large in the field. It was only by the complete 
cooperation of the Navy, its contractors, the Bureau 
of Standards, and the contractor of the Division that 
the project reached successful combat use. 

19 SERVO LINKS 10 

The use of full-span elevons results in larger hinge 
moments than would ensue from the use of airplane¬ 
like control surfaces, elevators, and a rudder. The 
requirement of constant angle of attack, which re¬ 
quired this design, was, however, deemed to be in¬ 
flexible; indeed, all experience of the Division indi¬ 
cates that its importance cannot be overestimated. 
A severe requirement was therefore placed on the de¬ 
signers of the servo link: to produce a mechanical 
power amplifier of very high gain with a minimum 
of phase lag. 

The initial attempt used a continuously running- 
electric motor. Four jaw clutches imparted , to the 
two elevons motion to increase or to decrease the lift 
of each, independently of any motion which might be 
taking place on the part of its mate. Thus, if the fol¬ 
lowing table shows the action of the four clutches: 

Increase Lift Decrease Lift 

Left Elevon Clutch No. 1 Clutch No. 2 

Right Elevon Clutch No. 3 Clutch No. 4 

an increase in rate of turn to the right would be exe¬ 
cuted by energizing No. 1 and No. 4, a decrease in 
glide-path angle by energizing No. 1 and No. 3, and 
so on. Flight tests showed that such a system imposed 
on the radio-controlling bombardier a task not likely 
to be successfully accomplished. Indeed, two flight 
tests resulted in such prompt failure that the decision 
was made immediately to interpose automatic flight- 
control equipment between the teledynamic signals 
and the servo link. (See Section 1.6.) 


The next design was successful. Here the continu¬ 
ously running motor with the four jaw clutches is re¬ 
tained, but the connection to the elevons is through 
a linkage so ingenious as to merit particular explana¬ 
tion. The four clutches now perform missile functions 
rather than elevon functions. In Figure 16 the four 
small pinions, R, L, U, and D, are driven through 
the clutches; R and L initiate the turning forces, and 
U and D the pitching forces. The whole linkage is 
supported by two heavy fixed members A . Suppose a 
“down” signal to be called for in the absence of any 
turning signal. The appropriate clutch will engage, 
bringing pinion D up to speed; this will drive the 
pitch sector say in a counterclockwise direction, and 
with it the pitch-input link. In the absence of any 
turning signal, the turn sector is stationary and holds 
rigid all the horizontally hatched system, including 
H r through E and F. Thus Hr, B, C, and the pitch- 
input link operate as a four-bar linkage, and the 
right-elevon output arm translates circularly, the 
stud moving counterclockwise about 0 4 . A con¬ 
necting-rod from the stud raises the right elevon. 
The rotation of the pitch-input link is transmitted 
through D to the internal jackshaft, which in turn 
gives through B' and C a counterclockwise rotation 
to the left-elevon output arm about 0 3 '. Now if the 
radii about Oi and 0 3 ' are equal and large compared 
with the arcs of travel, the motion of the elevons will 
be nearly identical. 

An exactly similar procedure covers turn. Suppose 
a “right” signal is called for, resulting in a clockwise 
rotation of the turn sector. Then the turn-input link, 
through C and B', gives a clockwise rotation to the 
external jackshaft, which in turn, through D, E, the 
bell crank Hr, B, and C, rotates the right-elevon 
output arm clockwise about 0 3 . Similarly, the stud 
on the left-elevon output arm will revolve about a 
suppressed axis 0/. Again, if the radii about 0 3 and 
Oi are equal and large compared with the arcs of 
travel, the motion of the elevons will be very closely 
symmetrical. 

Within the limits of linearity indicated above, 
these motions are subject to geometric superposition. 
In any event, departure from strict kinematic rigor 
is probably unimportant. Any rate of roll introduced 
by slight inequality of elevon action resulting from 
a pitch signal would be corrected by the automatic 
pilot. (See Section 1.6.) 

This scheme for a servo link was considered a tem¬ 
porary expedient to provide a means of flight-testing 
the glide bombs without delay. It was recognized 



26 


GLIDE BOMBS 


that the linkage was elaborate and not readily 
adapted to quantity production. Nevertheless, after 
a most exhaustive exploration of other forms of links 
involving hydraulic and pneumatic media of trans¬ 
mission, the linkage just described remained the 
basic method. The other designs were tested on the 
roll hunt table (see Section 1.5) and showed greater 
time lags resulting in more hunting, were deficient in 
output, or were too heavy for airborne equipment of 
this type. 

In the production version the sector-driven shafts 


were replaced by jackscrews, and the jaw clutches 
were replaced by positive-acting clutch gears similar 
to those used for controlling and quickly reversing 
the longitudinal feed on an engine lathe. These 
changes together with a general rearrangement of 
elements, produced a design which had somewhat 
better performance and was adapted to mass produc¬ 
tion. 

In another contract (see Chapter 7) the entire 
servo system was reviewed, and the automatic pilot 
was combined with the amplifying servo link. 


LEFT-ELEVON 
OUTPUT ARM 


TURN-INPUT 

LINK 


LEGEND 

VERTICAL = PITCH LINKAGE. 


HORIZONTAL* TURN LINKAGE 

CROSSED LINKAGE OUTPUT ARMS 


PITCH- 

INPUT 

LINK 



RIGHT-E LEVON 
OUTPUT ARM 


STUD 


Figure 16. Schematic drawing of elevon drive linkage. 








































































Chapter 2 

AZON AND RAZON 


21 INTRODUCTION 

T he glide bombs discussed in the preceding chap¬ 
ter have a lift-drag ratio of about seven. The long 
range implicit in this ratio, approximately 13 miles 
from an altitude of 10,000 feet, yields relative invul¬ 
nerability to antiaircraft opposition and is achieved 
by the use of large control and supporting surfaces. 
Thus increased range is obtained at the expense of 
ease of carriage. Aircraft carrying glide bombs are 
limited in their load to two or, at most, three missiles. 
The aircraft requires modifications to permit the car¬ 
riage of the missiles; carrying them externally im¬ 
pairs in a measure the performance of the aircraft and 
clearly marks the airplane for a special attack by 
enemy pursuit ships. For this reason NDRC under¬ 
took the development of a guided missile which could 
be carried within standard bombardment aircraft, be 
hung on standard bomb racks, be produced by modi¬ 
fying standard bombs, if possible, and which would 
require for its control the minimum of specialized 
equipment in the carrying plane. 

22 GENERAL 

In order to achieve this objective it was early real¬ 
ized that the bomb must provide the major portion of 
the lift and that steering forces must be provided by 
the bomb casing itself through yawing and/or pitch¬ 
ing it in the wind stream. A well-established agree¬ 
ment between the Army Ordnance and the Army Air 
Forces had decreed that bombs carried within air¬ 
craft should have no protuberances which extend 
beyond the smallest square prism which circumscribes 
the body of the bomb. The designers of aircraft for 
the AAF, on the one hand, designed all bomb racks 
so that bombs fulfilling this requirement could be 
carried safely; Ordnance, on the other hand, under¬ 
took to see that all bombs were built to this specifica¬ 
tion. 

Within these bounds each group of designers had 
flexibility, but the limits were held rigid. The Divi¬ 
sion accepted these limitations and only after ex¬ 
haustive study within them attempted to secure a 
relaxation at the expense of reducing the bomb load. 
The implications in this restriction are powerful. 


Any wings or supporting surfaces may have a span 
not exceeding the diameter of the bomb casing by a 
factor greater than the square root of two. In order to 
have an appreciable area their chord must be great— 
several times their span—so that their aspect ratio is 
much smaller than normal. The provision of a long 
lift surface forward of the center of gravity impairs 
the weathercock stability. It was finally found that 
the gain in lift due to such “wings” was so small in 
comparison with their cost in stability at trim that 
they were abandoned, the bomb body itself develop¬ 
ing the major portion of the lift. 

This simplification did not, however, produce a 
solution which was obviously of great promise. A. 
bomb casing in itself is not an aerodynamic structure 
of highly desirable profile. It is a form that would be 
no more than tolerable to the designer of a dirigible air¬ 
ship, and such aircraft do not contemplate speeds of 
550 mph—approximately the impact velocity of a 
dirigible high-angle bomb from 15,000 ft. 

The basic art applicable to the problem was mea¬ 
ger. Dryden 1 had made studies at the National Bureau 
of Standards on the fundamental aerodynamic prop¬ 
erties of standard bombs. Beyond this work, how¬ 
ever, neither the Division nor its contractors found 
much evidence of quantitative work which would 
indicate a probability of making standard bombs diri¬ 
gible. There had been considerable speculation, but 
the most informed opinion from aerodynamic advi¬ 
sers was definitely negative. The question of ade¬ 
quacy of lift from such a wing as a bomb casing has 
already been mentioned. The problem of exercising 
adequate control during the thirty-odd seconds of 
flight seemed difficult. The problem of determining 
what control to apply, granting the physical possibil¬ 
ity of applying it, seemed insurmountable. 

It may well be questioned why the Division under¬ 
took the project in the face of such negative advice. 
The desirability of converting near misses into hits— 
of compressing to within the dimensions of the actual 
target the relatively large number of craters that 
cluster about the center of impact with normal dis¬ 
tribution—seemed of great importance to the Division 
chief. However, had he had access to data on actual 
bombing errors and known that it was necessary to 
get the amount of control that finally proved feasi- 


27 


28 


AZON AND RAZON 


ble, it is doubtful if he would have recommended 
carrying the project through. 

In general there were two approaches to the prob¬ 
lem of controlling the bomb. In the first method the 
bomb was made to home on the target. This was 
planned with television and manual steering through 
a radio link or with an automatic homing device. The 
automatic homing devices which were considered uti¬ 
lized (1) light actuating a photosensitive system, (2) 
heat actuating a thermosensitive system, and (3) ra¬ 
dar homing, actuated by microwave reflections from 
the target. It was realized that a homing system must 
operate with its axis of scan parallel to the tangent to 
the flight path. Consequently, the use of pivoted 



Figure 1. Production models of Azon and Razon. 

vanes to align the television camera or the automatic 
homing receptor was contemplated. 

The second method contemplated a direct-sight 
method of control. In this method the bomb was to 
be controlled along a trajectory such that it continu¬ 
ously eclipsed the target from the point of view of the 
controlling bombardier. 

The first approach involves much more complicated 
control equipment. The second implies, as will be 
seen in Section 2.3, a much wider departure from the 
conventional parabolic trajectory. Each method found 
embodiment in a weapon standardized for combat. 
Felix, the automatic heat-seeking high-angle bomb, 
is discussed in Chapter 3. Azon, a high-angle bomb 


visually guided in azimuth only, reached combat in 
the Mediterranean, European, and Burma Theaters 
of operation. Razon, its two-coordinate counterpart, 
was in production for use in the Asiatic Theater as the 
war closed. Production models of Azon and Razon 
are shown in Figure 1. 

23 TRAJECTORIES 

Controlled high-angle bombs have two types of 
trajectory. For the homing type, the path is closely 
rectilinear when the target is motionless in the air 
mass; when there is target motion with relation to the 
air mass, the path becomes a pursuit curve. This type 
of trajectory is treated in Chapter 3. 

For visually guided bombs, the trajectory should 
be analyzed from two points of view. In the azimuth 
component, the trajectory does not usually require 
much curvature to effect the necessary corrections, 
and the problem of parallax, especially if the target 
is long in the direction of the bomb run, is not acute. 
It is the problem of parallax in the range sense which 
recommends the eclipse technique. If the controlling 
bombardier could estimate continuously the distance 
from the bomb to the ground and its first few time 
derivatives, he could doubtless be trained to steer a 
Razon in both coordinates as readily as an Azon in 
one. If, however, he keeps the bomb and target con¬ 
tinuously colli near with his own point of view, he is 
assured of a hit even if he has no idea of the time of 
impact. Eventually, the bomb will reach the ground; 
if it is continuously in the bombardier’s line of 
sight to the target, it will be on the target at impact. 

If this principle is adhered to throughout the flight 
of the bomb, the release will take place exactly at the 
target’s zenith since in the first instant the bomb is 
directly beneath the airplane. With a drop from 
20,000 ft and a speed of 150 mph, if collinearity is to 
be delayed for approximately 10 seconds, the drop 
should occur 2,000 ft before the target is directly be¬ 
neath the airplane. If collinearity is to be delayed for 
approximately 24 seconds, the drop should occur 
4,000 ft before the target is reached. In the case of the 
drop with the 10-second collinearity, some 30 seconds 
of control time would remain; for 24-second collinear¬ 
ity, about 16 seconds would be available for steering. 
A normal drop for an uncontrolled bomb would take 
place about 7,000 ft in advance of the target. 

Thus, in the eclipse method of range control all the 
forward velocity of the bomb must be killed, and 
during a goodly portion of its flight the bomb actu- 



ROLL STABILIZATION 


29 


ally regresses in the range sense. Figure 2 shows the 
three trajectories discussed in the preceding para¬ 
graph as projected on the vertical plane containing 
the path of the aircraft. It also shows the trajectory 
of a perfectly aimed, unguided bomb (Curve A). 

The increase in curvature of the controlled trajec¬ 
tories, as compared with that of a standard bomb, is 
a measure of the amount of control force which has to 
be applied. This technique, therefore, requires con¬ 
siderable maneuverability of the missile—not an 
ideal application for an airframe constrained to fit 
within the smallest square prism circumscribing the 
body of the bomb. Furthermore, it is undesirable to 
have the release point for guided bombs materially 
different from that of standard bombs because the 
miss, if control is lost, would be gross. 

Several methods have been suggested for combat¬ 
ing the problem of excessive curvature in the trajec¬ 
tory resulting from the eclipse technique. One of 
them (used by the Germans in Hs 293, a visually 
guided glide bomb) is to have a rocket propel the 
missile out in front of the aircraft so that collinearity 
can be established early without excessive trajectory 
warping. A second, tried by the Division’s contrac¬ 
tors, 2 involved decelerating the airplane so that it 
would drop behind the missile. This was accomplished 
by putting the bomber into a climb with minimum 
throttle. In order to get a retardation promising suc¬ 
cess, it was necessary nearly to stall the ship out. 
This method was used by the Germans in guiding 
FX-1500, a visually guided high-angle bomb. Soon 
after their success in sinking the Roma, the use of 
FX-1500 was abandoned when all the aircraft equip¬ 
ped to carry and control it were destroyed in a mass 
bombardment. Little data on the efficacy of this 
technique is, therefore, available. 

The Crab modification to the standard Norden 
bombsight provided a good solution. The problem 
was placed before the Fire-Control Division of NDRC 
by the Director of OSRD. The solution is discussed 
more fully in Section 2.6 and at length in the Sum¬ 
mary Technical Reports of Division 7, NDRC. The 
assumption was made that the time of flight of a 
Razon could be estimated. Then a partially silvered 
mirror was inserted in the optical train of the bomb- 
sight. The image of the falling bomb is projected into 
the field of view so that it appears superposed on the 
terrain at the point it would land if no further control 
were added. If the drop is perfect, the bombardier 
sees the bomb exactly on the target and makes no 
correction. Thus the departure from a normal trajec¬ 


tory is minimized. The residual range error after per¬ 
fect steering is equal to the error in estimating the 
time of fall multiplied by the forward velocity of the 
bomb at impact. 

2 4 ROLL STABILIZATION 

At the outset of the project two modes of control 
were proposed, each of which implied its own peculiar 
type of roll stabilization. One system, which has been 
uniformly attractive to newcomers in the field, in¬ 
volves a system of cylindrical coordinates. The axis 
of the system is tangent to the trajectory. In this 
system no attempt is made to prevent roll, only to 
restrict its rate to a reasonably low value, say 0.5 
revolution per second. Only one set of control sur¬ 
faces is provided for steering. Ailerons for restricting 
and controlling the rate of roll are also provided, 
usually in a plane normal to that of the steering sur¬ 
faces. The method is to roll the bomb so that a plane 
normal to the plane of the steering surfaces contains 
the target, the bomb, and the aircraft. The bomb is 
then given an angle of attack to produce a resultant 
of lift and gravity and, consequently, an acceleration 
in the direction in which correction of the trajectory 
is desired. 

Viewed from a fixed frame of reference in the bomb 
structure, lift can be produced only in a single plane, 
normal to the axis of roll and to the steering surfaces. 
If a turn to the right is required, the bomb is rolled 
until the steering surfaces are vertical and then the 
bomb is yawed. If a dive is called for, the bomb is 
rolled until the steering surfaces take the position of 
conventional elevators and then the bomb is pitched. 

In the second system the location of the three prin¬ 
cipal axes of motion remains unchanged, but the 
bomb is closely coupled to them. Two steering sur¬ 
faces are provided: the rudder operates in the plane 
of the roll and yaw axes, the elevator in the plane of 
the roll and pitch axes. This system, therefore, oper¬ 
ates in simple Cartesian space. 

The cylindrical-coordinate control has the advan¬ 
tage of not requiring absolute roll stability. Only a 
rate gyro is required which can be made sufficiently 
sensitive to hold the roll velocity to any desired value. 
At the outset of the guided-missile program free 
gyros, which are absolutely essential if the Cartesian 
system of control is to be employed, were difficult to 
procure. Further, a wind-driven or unpowered free 
gyro with two frames is sure to tumble if the bomb 
is pitched through 90 degrees and then yawed through 



30 


AZON AND RAZON 


HORIZONTAL POSITION OF AIRCRAFT /? SECONDS AFTER RELEASE 

® ® ® ® @V 

@ ® ® ® <©c 



76S43Z I Ol 234567 
SHORT OVER 

RANGE FROM TARGET - IOOO FT 


Figure 2. Profiles of eclipse trajectories compared with perfect uncontrolled bomb. 


















































ROLL STABILIZATION 


31 


90 degrees. Perhaps this property of the free gyro 
could be circumvented by providing more than two 
gimbal frames, but even if it were, all sense of right- 
side-upness would be lost after such a maneuver as 
the one just described. After a 90-degree dive the 
“top” of the bomb would point forward. A succeed¬ 
ing turn to the right would place the rudders hori¬ 
zontal with the “top” of the bomb on its left. 

The summation of these advantages led the inves¬ 
tigators strongly toward the cylindrical-coordinate 
system of control. Not only MIT and Gulf Research 
and Development favored it for the high-angle bomb, 
but Douglas strongly favored it for Roc (see Chapter 
4), and Polaroid for Dove Eye (see Chapter 3). It 
was also strongly urged for the Jeffries 3 bomb under 
development by the British. It has been treated at 
length in this report as an example of exploration 
which the Division found unprofitable. The Division’s 
experience all points to the conclusion that it is much 
better to plan trajectories which will not involve succes¬ 
sive turns approaching 90 degrees about the axes of yaw 
and pitch than to try to provide a stabilization system 
which will tolerate them. 

2 41 Origin and Nature of Roll Torques 

Roll torques occur in all bombs. They are deliber¬ 
ately provided in some German bombs by designing 
in a screw asymmetry to arm the fuze and in the 
British Tallboy (a 12,000-lb, deep-penetration bomb) 
to improve stability about the yaw and pitch axes. 
In bombs manufactured in the U. S., it is never pur¬ 
posely provided, but the difficulty of holding manu¬ 
facturing tolerances within limits which will elimi¬ 
nate casual screw asymmetries is so great that in 
general all standard bombs in flight have a random 
roll velocity. 

Induced Roll Torques. Induced roll torques arise in 
dirigible bombs from the application of control about 
the axes of yaw and pitch. Such induced roll torques 
occur with conventional control structures when yaw 
and pitch are applied simultaneously. They appear to 
arise from three sources: 

Consider a bomb which has been pitched. Such a 
bomb has an angle of attack at the elevators. If it 
now be yawed, one elevator will lose lift because of 
the reduced speed of the elevator on the inside of the 
turn. This torque is probably small and should be 
transient in nature, occurring only during the period 
while the bomb is approaching its trim attitude. All 
rotation thereafter is at an axis so remote, certainly not 


less than 2 miles, that any differential velocity in the 
lift surfaces is wholly negligible. This property can be 
disclosed only by transient studies in the wind tunnel. 



ANGLE OF ROLL 0 IN DEGREES 


Figure 3. Roll torques in cruciform bomb as function 
of roll angle. 

A flat plate in a wind stream tends to orient itself 
so that it is perpendicular to the direction of flow. So 
long as the plate is parallel to the wind stream, it is 
in a state of unstable equilibrium and but little force 
is required to hold it there. If a bomb is pitched with¬ 
out yaw the plane of the rudders is parallel to the 
wind stream and only slight aileron power is required 
to maintain the roll orientation. If a pitched bomb 
then be yawed, there is a torque as both the rudder 
and elevator attempt to attain a position normal to 
the wind stream. This torque disappears when a roll 
attitude is reached which places the structure 45 
degrees away from the desired orientation. 
























































32 


AZON AND RAZON 




NO YAW OR 
PITCH 


PITCH WITHOUT 
YAW 


PITCH AND 
YAW 


5 


< 

UJ 



UJ 

(T 


a 


Figure 4. Drawings of cruciform bomb in attitudes of 
no yaw or pitch, pitch, pitch and yaw. 


Figure 3 shows the variation of roll torque with roll 
attitude as measured by MIT in the wind tunnel. In 
these curves the unit of roll torque is the maximum 
restoring torque that could be developed by the aile¬ 
rons planned. The slope of the curve at zero roll, as 
well as the value of roll torque at that orientation, is 
of crucial importance, a negative slope indicating sta¬ 
bility and a positive slope instability. Thus Curve A 
(pitch only) and Curve B (yaw only) show roll stabil¬ 
ity; Curve C (simultaneous roll and pitch), however, 
shows marked instability. 

The third source of induced roll torque arises from 
shadowing of the control surfaces by the body of the 
bomb. Figure 4A shows a bomb without yaw or pitch. 
The rudders a, a f and the elevators b, b' are parallel 
to the wind stream which is in the plane of the draw¬ 
ing. They therefore produce no torque. In Figure 4B 
the bomb has been pitched so that the lower rudder 
a' is shaded by the bomb casing. Since the rudders 
are parallel to the wind stream, no roll torque is 
developed in spite of the asymmetrical exposure of 
the rudders to the wind. In Figure 4C both yaw and 
pitch are applied. Now in addition to the lower rud¬ 
der a' the remote elevator b' is shaded by the bomb 
body. Further, both the rudder and elevator planes 
are inclined to the wind stream so that the asym¬ 
metrical exposure of each results in a roll torque. 

The torques developed by shading are in opposi¬ 


tion. The excess exposure to the wind of the upper 
rudder produces a counterclockwise moment; that 
from the exposure of the elevator in the foreground 
produces a counterclockwise moment. Figure 5 is a 
contour map of the roll torques produced by simul¬ 
taneous yaw and pitch. As might be expected from 
the preceding two paragraphs, two pairs of orthog¬ 
onal axes define the attitudes producing zero roll 
torque. Displacement in pitch or yaw alone results in 
stable equilibrium. Equal displacement in yaw or 
pitch results in zero roll torque, but the equilibrium 
is unstable. 

While the torques measured in the wind tunnel 
show requirements three times in excess of those 
planned, this is not the whole story. Flight tests 
showed that torques at least three times greater than 
those measured could be developed, because of auto¬ 
gyration once a roll velocity is established. 

2 4 2 Reduction of Roll Torques 

The discussion in Section 2.4.1 as to the origin and 
nature of roll torques points clearly to the method of 
their minimization. It is to produce a structure which 
is symmetrical about any plane passing through the 
axis of roll. A dirigible high-angle bomb having such 
a configuration was designed by MIT 4 and is shown 
in Figure 6. Lift is obtained by the forward shroud 
which is placed so that its center of lift is substanti- 



Figure 5. Contour of roll torque as function of yaw 
and pitch. 




































































ROLL STABILIZATION 


33 


ally at the center of gravity of the missile. Stability 
about the yaw and pitch axes derives from the outer 
of the two empennage shrouds. The inner empennage 
shroud, capable of being rotated about axes parallel 
to those of yaw and pitch, serves as a rudder and 
elevator. 

This design was fully tested in the wind tunnel at 
MIT and was found to develop no roll torques at 
any combination of pitch and yaw attitudes. It was 
flight-tested at Eglin Field in February 1943 5 and 
developed no tendency to roll in spite of being vio¬ 
lently pitched and yawed during its flight. In addi¬ 
tion to exceeding the permissible compass for stand¬ 
ard stowage by a very wide margin, however, it had 
intolerably large values of hinge moment, with the 
corresponding implication of excessive power re¬ 
quirement. 



lift cylinder control cylinder 


Figure 6. MIT cylindrical-fin bomb. • 

A design to establish a working compromise be¬ 
tween minimum roll torque and tolerable hinge mo¬ 
ments was worked out by Gulf (Figure 7) which 
became the prototype for the Felix (high-angle heat¬ 
seeking bomb). (See Chapter 3.) This bomb proved 
completely stable in roll and had adequate maneuver¬ 
ability. 6 Power from small battery-powered motors 
was adequate to operate the control surfaces. 

A second method of minimizing roll torque lies, of 
course, in the limitation of the control of the bomb to 
motion about only one of the axes of yaw and pitch. 
From the point of view of roll alone it is immaterial 
which axis is selected. For the purposes of developing 
a useful weapon, problems of sighting necessitate the 
selection of azimuth or yaw as the single component 
of control. 

The problem of roll stability, then, confronted the 
Division with the choice between one-axis control 
and exceeding the limits which had been set on the 
overall size of the missile. Both solutions were ac¬ 


cepted. As the Azon development was pushed to 
completion, studies were made to determine how 
seriously extensions of portions of the control struc¬ 
ture beyond the smallest square prism circumscribing 
the body of the bomb would vitiate the military use¬ 
fulness of Razon and Felix. The compass of Azon was 
held within the limits originally established. In April 
1943, at the completion of the studies just described, 



STABILIZER 

Figure 7. Gulf octagonal-fin bomb. 

the maintenance of such limits for bombs with two- 
axis control appeared to the contractor to be hope¬ 
less, 3 and further efforts in that direction were not 
made. This opinion was not unanimously shared by 
the Division. The cam-controlled drop reproduced in 
Figure 13 (see Section 2.9) was successfully roll- 
stabilized with 10-degree deflection of the elevators 
and a simultaneous regime of rudder deflection rang¬ 
ing from 5 degrees right to 10 degrees left. Further 
studies in this field might well be rewarding. 

2 4 3 Gyro Developments 

The decision to abandon a cylindrical-coordinate 
control in favor of a Cartesian coordinate system 
(see Section 2.4) implies, as already shown, the re¬ 
quirement of a free gyro to control the ailerons. Such 
a gyro alone, however, can produce nonoscillatory 
roll stabilization only with difficulty. MIT therefore 
developed a system involving a free gyro to deter¬ 
mine the roll orientation and a rate gyro to damp the 
roll oscillations. 

Consider a free gyro with contacts on the outer 
gimbal frame to close two pairs of contacts with a 
dead band of ±<f >o between them. Thus one contact 
was closed if <f> > + <£ 0 ; if <t> < — </>o, the other con¬ 
tact is energized. Similarly, the rate gyro has two 
contacts separated by a dead band of ± <po. If 

a See Chapter 13 for further comments by the contractor’s 
project director. 







34 


AZON AND RAZON 


0 > + 0o, one set of contacts closes; if <j> < — <j> o, 
the second set of contacts closes. The contacts on 
the free and rate gyros are interconnected through 
relays to give aileron action in accordance with the 
following schedule: 


4 > < -0o, 0 < - 0 o CW b 

0 < -00, - 00 < 0 < 00 CW 
0 < — 00, 0 > 00 No 

-00 < 0 < 00 , 0 < -00 CW 

— 00 < 0 < 0o, — 00 < 0 < 0o No 

— 00 <0 < 00, 0 > 00 CCW 

0 > 00, 0 < — 00 NO 

0 > 00 , - 00 < 0 < 00 CCW 

0 > 00 , 0 > 0 CCW 


aileron action 
aileron action 
aileron action 
aileron action 
aileron action 
aileron action 
aileron action 
aileron action 
aileron action 


As the program advanced, this basic conception of 
a free and a rate gyro was maintained, although the 


b CW (Clockwise), CCW (Counterclockwise). 


contact arrangements were altered to give different 
schedules of aileron action. The initial gyros were 
air driven. The wheels were enclosed in a sealed cover 
which was connected with the vacuum line in the 
airplane. Jets to drive the wheels drew air from the 
surroundings through vents in the covers. 

Figure 8 shows the motion of the free gyro in ref¬ 
erence to the bomb structure as an Azon is released 
and falls. The Azon is carried in standard bomb 
racks with the rudder and elevator fins inclined to 
the vertical by an angle of 45 degrees. The position 
of the bomb and of the free-gyro frames for this con¬ 
dition is shown in Figure 8A. As the bomb falls it 
must be rolled so that elevator and rudder become 
parallel to the desired orientation of the pitch and 
yaw axes. Simultaneously, the bomb noses down as 
it proceeds along its parabolic trajectory. This 
change, together with the changing orientation of the 
gyro, is shown in Figures 8B,8C, and 8D. Figures 8E 



Figure 8. Action of free gyro during Azon drop. 









LIFT AND MANEUVERABILITY 


35 


and 8F show relationship of the gyro gimbal as the 
bomb is yawed to the right and to the left. 

In its final form, the gyro assembly was electrically 
driven from the 24-v battery in the bomb tail. The 
dead band was removed from the free-gyro contacts, 
which were fnade movable and coupled to the rate 
gyro so that the position of the free-gyro contacts, 
which control the ailerons, is biased by the rate of 
roll. Under this arrangement there is no neutral; that 
is, the ailerons are always operating to produce 
either a CW or a CCW roll torque. The schedule now 
becomes much more simpty stated: 

(<f> + k<j>) <0 CW aileron action 
(</> + k<j>) > 0 CCW aileron action 

The factor k must have the dimensions of time and is 
set at 0.5 second. In addition an overriding rate-of- 
roll contact is provided which limits the roll velocity 
to + # max , which was arbitrarily set at ± 36 degrees 
per second. 

The roll motion of the bomb can now be defined. 
If the displacement from correct roll orientation is 
in excess of 18 degrees, the bomb will return to 
(f> = +18 degrees at an angular velocity of —36 
degrees per second. From this position it will con¬ 
tinue along the regime 

4> = lSe~ 2t 

The foregoing idealizes the action of the ailerons 
by assuming absence of time lag in their operation 
as (<f> ± 0.5 4>) changes sign. This time lag is present 
largely on account of the time constant of the sole¬ 
noids which operate the ailerons. It introduces a roll 
hunt of approximately 2J4 c. The amplitude of the 
hunt is proportional to the square of velocity and 
reaches approximately 3 degrees at the end of a 
15,000-ft drop. 


LIFT AND MANEUVERABILITY 


It was explained in Section 2.3 that the attainable 
trajectory of a dirigible bomb is determined by the 
transverse lift which it can be made to achieve. In¬ 
deed, in the steady state the performance of such a 
missile is completely determined by two dimension¬ 
less coefficients: the lift coefficient Cl and the drag 
coefficient Cd. These are the well-known coefficients 
of the aerodynamicist and are defined by: 

Lift , ^ Drag 


C L = 


hpV*A 


and Cd = 


\ p V 2 A 


( 1 ) 


where p is the air density, V the velocity of the mis¬ 
sile in the air mass, and A is an area function. In this 
chapter A is the cross-sectional area of the bomb 
body—a confusing choice to the aircraft designer, 
who usually uses this symbol to denote wing area, 
but convenient here in that it makes the various 
values of Cl and Cd directly comparable, irrespective 
of the areas of the control and lift surfaces which the 
course of the research brought under examination. 

The drag coefficient is not of great significance in 
the visually guided bomb. It limits the terminal ve¬ 
locity and increases the trail angle, 0 thus influencing 
the sighting problem in the range sense. 

The total lift attainable divided by the mass of 
the bomb gives the maximum attainable transverse 
acceleration. It is more usual to divide by the weight, 
however, and express the acceleration in “times 
gravity” (g’s). 

A , + . iC L 'pV 2 A . , 

Acceleration = -^-m g s (2) 

where Cl is the maximum attainable value of C l . 
This is a convenient measure of the maneuverability 
of a missile, but it is tied up with the air density and 
the velocity. A still more useful measure is the mini¬ 
mum turning radius. A radius of curvature, R, re¬ 
quires a centripetal acceleration. 

V 2 

Centripetal acceleration = (3) 


Setting equation (2) equal to equation (3) and solv¬ 
ing for R: 


W 

hpC L 'Ag 


(4) 


The acceleration due to gravity g is inserted to pre¬ 
serve the dimensional integrity. This parameter is 
still a function of p and it is usually given as a sea- 
level value. 

The most useful parameter which could be estab¬ 
lished to define the performance of a visually guided 
bomb of the Azon or Razon type would be the cor¬ 
rigible error. This is a “judgment” figure however, 
involving an estimate of the skill of the bombardier 
in recognizing the sense of his error and making the 
appropriate correction. A convention could be estab¬ 
lished defining the “correction coefficient” of a guided 


c Trail angle is the angle between the line of sight from the 
aircraft to the bomb and a vertical through the aircraft. This 
definition assumes constant rectilinear flight by the aircraft 
after release and neglects changes in air density. 







36 


AZON AND RAZON 



PERFORMANCE 
WITH FULL RUDDER 

SEA-LEVEL AIR 
VELOCITY 800 FPS 

MAX.LIFT MIN. RADIUS 
1.33 G 15pOO FT 


MIT DESIGN MEDIUM FINS 



0 2 4 6 8 1012141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 



MIT DESIGN 


MAX.LIFT 
0.74 G 


1.3 

1.2 

o ' 11 ' 1 
v— 1.0 
5 0.9 
o 0.8 
It 0.7 

MIN.RADIUS g o.6 
27,000 FT 3 o.5 

t 0.4 
□ 0.3 
0.2 
0.1 
0 



SHORT FINS 


0 2 4 6 8 10 12141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 

























































































































0 2 4 6 8 10 12 141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 



MAX.LIFT MIN.RADIUS 
0.83 G 24,000 FT 


PRODUCTION AZON 



0 2 4 6 8 101214 1618 
TRrM ANGLE OF ATTACK 
IN DEGREES 


Figure 9. Developmental evolution of Azon and Razon. 







































































































LIFT AND MANEUVERABILITY 


37 



LIFT CYLINDER CONTROL CYLINDER 


PERFORMANCE 
WITH FULL RUDDER 

SEA-LEVEL AIR 
VELOCITY 800 FPS 

MAX.LIFT MIN. RADIUS 
1.27G 15,700 FT 


MIT CYLINDRICAL FIN 



0 2 4 6 8 1012141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 



GULF OCTAGONAL FIN 



0 2 4 6 8 1012 141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 



FLARE J 

RAZON MK1 


MAX.LIFT 
0.67 6 


MIN. RADIUS 
29,800 FT 



TRIM ANGLE OF ATTACK 
IN DEGREES 



RAZON MK 4 




MAX. LIFT 
0.97 G 


MIN. RADIUS 
20,500 FT 



0 2 4 6 81012141618 
TRIM ANGLE OF ATTACK 
IN DEGREES 


Figure 9. ( Continued ). 
































































































38 


AZON AND RAZON 


bomb as the deflection of the point of impact due to 
full control in the last n seconds in a drop from h 
feet. The computation of such a coefficient is, how¬ 
ever, transcendental and requires the use either of 
step-by-step integration or some such aid as the 
Rockefeller differential analyzer. The MIT group 
working on this project and the associated group 
working on Felix used point-by-point computations 
until the method was established. Thereafter, they 
used the differential analyzer most effectively. 

Figure 9 shows several of the models of dirigible 
bombs that were tested as the program advanced 
and the various parameters which define their 
maneuverability. 

2 6 SIGHTING AND PARALLAX* 

Section 2.3 explains the ruse by which the problem 
of parallax in range steering is circumvented. Sev¬ 
eral descriptions of the Crab attachment to the Mark 
15 bombsight have been written. That of the Gulf 
Research and Development Company 7 is brief and 
clear. It is quoted in its entirety: 

In a standard uncontrolled bomb, the angle subtended by 
the target and bomb with the airplane e closes to zero at impact 
time at an almost linear rate. Thus, the fundamental require¬ 
ment of the sighting device is that it provide a means by which 
the angle subtended by the bomb and target can be compared 
continuously with one which closes to zero at impact time at 
a constant rate. 

In the Norden bombsight a mechanism exists which pro¬ 
vides just such a motion in the tracking mirror which closes 
the angle subtended by it at the release of the bomb and the 
Trail Angle of the bomb at impact in exactly the time of fall 
of the bomb as put into the bombsight through the Disc Speed 
setting. The details of the mechanism by which this result is 
obtained are too involved to describe here; a description can 
be found in the technical literature describing the sight. 

Two adaptations of the Norden bombsight were built for 
utilizing this mirror motion, one by the contractor involving a 
double cross-hair method, and another by the Franklin Insti¬ 
tute involving a double mirror arrangement similar to that 
used in a ship’s sextant. In the contractor’s version, an auxil¬ 
iary cross-hair at the eyepiece reticle is moved across the field 
by the regular sight mechanism in a manner such that at the 
end of the assumed time of fall the moving and fixed cross¬ 
hairs of the sight are coincident. Hence if, by steering, the 
bomb is held on the moving fiducial, and, by precessing the 


d See also Aiming Controls in Aerial Ordnance , G. A. Phil- 
brick, Vol. 3, Part I, of the STR of Division 7, NDRC, and the 
15th Bi-Monthly Report to NDRC of Division 5. 

e More exactly the quantity (tan 0 + tan <t>) closes to zero 
at a linear rate, where 0 is the angle subtended by a line to the 
target with the vertical, and <£ is the angle subtended by a line 
to the bomb with the vertical, also known as the trail angle. 


vertical gyro, the fixed fiducial is held exactly on the target, 
then the proper steering criterion is attained. 

This arrangement suffers from limitations in the field of view 
of the Norden telescope, which, even after permissible modifi¬ 
cations, allows steering during only about the last 12 seconds 
of flight. Furthermore, if the synchronization of the cross-hair 
on the target is not perfect, then the bombardier has the addi¬ 
tional worry of having to continually precess the gyro in order 
to hold the fixed cross-hair on the target. 

Both of these objections are eliminated in the adaptation 
worked out by the Franklin Institute/ Philadelphia, Pennsyl¬ 
vania, in which there are split lines of sight, one to the bomb 
and one to the target. If the angle between the two lines of 
sight is caused to converge at the proper rate, reaching zero 
angle (i.e., coincidence) at the end of the assumed time of fall, 
a perfect drop would show the bomb eclipsing the target 
throughout its flight. In the case of an imperfect drop, any de¬ 
viations of the bomb from a perfect eclipse position can be cor¬ 
rected by control applications. Analysis will show that, except 
for errors in predicting the time of fall and errors in the mech¬ 
anism reproducing this time, all normal sighting errors (e.g., 
improper range synchronization or leveling errors) can be 
eliminated by steering. 

The Franklin Institute adaptation was named the Crab 
sight (sometimes called Crab No. 1) and was used in all drop 
tests. For the line of sight containing the target, the normal 
moving mirror of the sight is used, and it simply continues to 
track the target until the bomb hits. For the second line of 
sight, a small mirror is mounted on the telescope objective so 
as to reflect into view a line of sight slightly rearward of the 
vertical. The second line of sight will reflect the bomb image 
into the field, since in a normal drop the apparent position of 
the bomb is always slightly rearward of a true vertical line ex¬ 
tending from the airplane. The bombardier sees two fields of 
view, one including the area around the target, the other in¬ 
cluding the terrain directly below the airplane. 

Before the Crab sight was tested in flight, there was some 
concern about the possibility of the confusion which might 
exist in trying to identify the target and the bomb in the two 
overlapping fields of view. In practice, there is little difficulty. 
A red filter in front of the bomb mirror is of some value, but 
even more important is the fact that the scenery reflected into 
view by the small fixed mirror (called the Crab mirror) is 
always moving, although the flare itself is relatively motion¬ 
less. Thus, it is necessary merely to concentrate one’s attention 
onto the relatively stationary flare image and scenery around 
the target. The exact arrangement of parts can be seen in the 
photographs shown in Figure 10. 

For a technically more exact description of the Crab sight, 
reference is made to Figure 11. In this figure, the variation of 
the angular position of the target with respect to the vertical, 
0, and the Trail Angle, <J>, for the case of a perfectly dropped 
standard bomb is shown. Since the airplane is assumed to fly 
along an unaccelerated g path after release of the bomb, the 
value of tan 0 must obviously decrease at a linear rate. The 
variation of tan <£ can be found in making a complete trajec- 


f Operating under contract with Section 7.2, NDRC. 

g In practice this is difficult to accomplish precisely because 
of the sudden change in weight of the airplane which occurs 
when a bomb is released. 





SIGHTING AND PARALLAX 


39 





Figure 10. Crab sight. Crab attachment in place on Mark 15 bombsight. 










40 


AZON AND RAZON 


tory calculation, which shows that, in general, for a bomb 
dropped from 15,000 ft. altitude the Trail Angle will start out 
at release with a value approximately two-thirds its final value, 
toward which it gradually increases during the flight. This rate 
of variation of the Trail Angle is shown in Figure 11. 

In the lower curves shown in Figure 11, the position of the 
target and of the bomb in the field of the bombsight telescope 
are plotted. Since for the case shown it is assumed that the 
synchronization is perfect, the target appears squarely in the 


PATH OF AIRPLANE 



IMPACT 


150 

100 

50 

0 

-50 





-1-1-i 

POSITION OF TARGET 

1 



3 _ 

BOMB 

WITHOUT 


AND 

BOMB IN FIELD 



o 

2 

PARALLAX 


OF BOMBSIGHT TELESCOPE 


_ y 

0 


r 


=1 

1- 




U- 

<£ 




-r?— 

TARGET 




K 

c 

cr _ 










Q- 

CL 

< 

—L 

BOMB WITH 
PARALLAX 








10 15 20 25 30 

TIME AFTER RELEASE IN SECONDS 


Figure 11. Angular relationship in Crab-sight use. 


center of the field throughout the flight. The position of the 
flare varies, however, for two reasons: first, because of the 
variation in Trail Angle, and, second, because of parallax which 
results from the fact that the bombsight is not mounted di¬ 
rectly over the bomb bay but a considerable distance ahead of 
it (24 ft. in a B-17). Thus, the bomb appears to have a much 
larger Trail Angle early in flight, as is indicated in Figure 11. 
Near the end of flight this variation in position of the flare 
image is negligible, so that for all practical purposes the bomb- 
sight will indicate no error for a perfect drop. 

The residual errors, other than those of steering 
in Razon, are thus due wholly to errors in estimat¬ 
ing the time of fall. This time, in addition to being 
a function of altitude of the release and of the target, 
is also dependent upon the total amount of steering 
to which the bomb is subjected during its fall. A 


large number of drops were made to establish em¬ 
pirically the relationship between time of fall and 
steering (Figure 12). It was found 8 that steering in 
the direction to decrease range, i.e., “down” elevator 
introduces no increase in time of fall. This is because 
the warping of the trajectory more nearly into line 
with gravity counteracts the increase in drag due to 
rudder deflection. If, then, At is the increase in time 
of flight due to steering, 2(4) is the totalized time 
of “up” elevator, and 2(4) is the totalized time of 
rudder application: 

A* = a2(4) + 62(4) (5) 

where a and 6 are the empirically determined con¬ 
stants. 

Jag (Just Another Gadget) is a device developed 
to insert this correction into the bombsight. To a 
reasonable approximation it is accomplished by sub¬ 
tracting from the tracking mirror’s angular velocity 
an incremental velocity proportional to At. The prob¬ 
able error of Razon in range without Jag 8 is estimated 
at 10 mils. Jag will reduce this by about two-thirds. 


27 RADIO 

The problem of providing a suitable radio link for 
the remote control of missiles was a major activity 
of the Division and is covered comprehensively in 
Chapter 6. 

The initial work on the Azon-Razon program used 
a radio having two r-f carriers. These were pulsed by 
a commutator keying circuit. The ratio of time-on to 
time-off of the pulse on one carrier was determined 
by the position of a rudder-control stick. An in¬ 
tegrating circuit following the detector in the re¬ 
ceiver caused the servo link to make the rudders as¬ 
sume a position corresponding to that of the rudder- 
control stick. A similar chain produced proportional 
control of the elevators. The r-f section of the re¬ 
ceivers were superregenerative for compactness; the 
gain was adjusted so that the final stage saturated 
with the lowest expected signal. 

The field tests at Eglin Field in April 1943 showed 
that although proportional control was available it 
was not used; that is, the controlling bombardier put 
the rudders hard over, adjusting the amount of error 
correction by the timing of full rudder commands. 
(See also Chapter 10 for remarks about on-off versus 
proportional control and simulative methods of their 
analysis.) Accordingly, the standard RC-186 trans- 

















































RADIO 


41 




Figure 12. View of target and Razon through Mark 15 bombsight equipped with Crab attachment. 


























42 


AZON AND RAZON 


mitter having six audio tones was employed on all 
further work with an on-off modulation—two tones 
with Azon, four tones with Razon. 

A super-regenerative receiver with a resistance- 
capacitance filter in the audio end of the circuit 
operated two relays giving “right” and “left” signals 
to the rudders. 

This receiver, with minor changes for production, 
went into combat as standard equipment for Azons. 
It was considered deficient in selectivity and stability 
and was suspected of being subject to false operation 
from microphonics occurring in the Azon tail. The 
Army had, however, instituted procurement and 
they were available. 

In order to obtain a more reliable radio for Razon, 
the Division, through Contract OEMsr-240 with 
MIT, procured the design of a sharply tuned, crystal- 
controlled superheterodyne receiver with a tuned 
transformer for audio selection. Some forty of these 
receivers were used in the Razon development pro¬ 
gram. After the initial difficulties were overcome, 
they were fully reliable. Other activities of the Divi¬ 
sion in developing a suitable radio for control are 
discussed in Chapter 4 and in Chapter 6, which is 
concerned with the radio problem in general. 

28 ACCESSORY COMPONENTS 

281 Servo Links 

Since the Azon and Razon are manually controlled, 
the servo-link problem is a simple one. This is not to 
say that all missiles can always be manually con¬ 
trolled through simple servo links. In this case, it 
developed that the dynamics of the servo loop had 
phase-gain relationships such that a human operator 
was able to cope with its response readily without 
introducing serious hunting. 

This was not fortuitously achieved. An engineer¬ 
ing sense was developed by Gulf which led them to 
choose bomb designs having only a moderate com¬ 
pliance and a reasonably low natural period about 
the pitch and yaw axes. Periods of 1 to 2 seconds 
were found to result in bombs relatively easy to 
steer. Lower periods obtained with higher compliance 
resulted in bombs which got out of control; higher 
periods produced bombs that “felt too stiff.” 

A simple 24-v geared motor, standard in the air¬ 
craft industry, formed a very satisfactory servo 
link. 


282 Flares 

In order to follow the bomb during its flight a 
pyrotechnic flare was mounted on the tail. Reliable 
ignition was an annoying but not a profoundly tech¬ 
nical problem. It was solved by using electric ignition 
of a time-delay powder train, which in turn fired the 
flare 8 seconds after the ignition circuit was energized 
by the bomb release. This is not wholly foolproof. 
One plane was burned up when a flare-equipped 
bomb dropped on the taxi strip before take-off; an¬ 
other was seriously burned in the air when a bomb 
hung up in the bomb bay. The flare-energizing circuit 
should be closed by a contact on the tail fuze, which 
depends for arming on travel for several hundred feet 
through the air. 

The use of colored flares and selective filters on the 
Crab sights for mass attacks was studied in coopera¬ 
tion with Division 16 and Division 11. Wesleyan 
University, under their contract with Division 11, 
experimented with various color formulas. 9 ’ 10 In gen¬ 
eral, the results were not too successful. Most pyro¬ 
technic color formulas are subtractive, as are filters. 
Thus, a bombardier using a red filter, for example, 
would hardly distinguish between a white flare and a 
red one. Actually, it was found that a bombardier 
had no trouble concentrating on his own bomb even 
in tight formations with all flares white. 

2 8 3 Fuze Arming 

The detonation of standard general-purpose bombs 
is accomplished typically by two fuzes, one in the 
nose and one in the tail. These fuzes are armed by 
wind vanes driven by the wind stream after the bomb 
has been released. An arming wire, which is with¬ 
drawn on release of the bomb, pins the wind vanes 
and prevents their rotation from drafts in the bomb 
bay. 

It is perfectly simple to apply the standard nose 
fuze to Azon and Razon. This was done. The tail 
fuze presented a more troublesome problem since the 
mechanism in the tails prevented the use of wind- 
stream arming. Electric arming motors were sug¬ 
gested and tried. It was conceivable that failures of 
the circuit would cause the bombs to arm before they 
were released and this system was, therefore, deemed 
unsafe. 

A cup-anemometer drive to be mounted on the 
side of the Azon or Razon tail was offered by the Air 
Technical Service Command. This was tested with 



FIELD TESTS 


43 


some misgivings since the trying experience of the 
Division with roll stabilization raised doubts as to 
the possible introduction of screw asymmetry in the 
structure with the anemometer on one side. Fortu¬ 
nately, these fears proved groundless, and this sys¬ 
tem was applied to both bombs. 

29 FIELD TESTS 

The program of developing a guided bomb finds its 
most active phases at the testing ranges. Much can 
be done in the laboratory; as we have seen, the wind 
tunnel is a most necessary tool in establishing the 
course along which an investigation may proceed. 
For the great bulk of the work, however, actual drop 
tests form the only technique so far established for 
developing and proving a design. 

For such work the Division relied almost entirely 
upon the military. In the program on glide bombs 
(see Chapter 1) the Navy Bureau of Ordnance fur¬ 
nished facilities. For the Azon and Razon program, 
the field testing was done at establishments of the 
Army Air Forces. It is impossible to say too much in 
appreciation of the cooperation rendered by the offi¬ 
cers of the Air Technical Service Command and its 
predecessor group, the Air Materiel Command. In 
spite of being short of aircraft and range facilities for 
the multiplicity of projects that they had in hand, 
they arranged to share what they had with the Divi¬ 
sion, sometimes at the expense of their own projects. 
It was not until December 1944 that suitable facil¬ 
ities with aircraft, ranges, ground transportation, and 
maintenance and shop facilities were established at 
Wendover, Utah, in which the Air Forces group 
charged with the development of guided missiles had 
a vested interest. Theretofore, they had worked— 
and the Division perforce with them—as more-or- 
less welcome guests of Ordnance Department at 
Aberdeen, Md., the AAF Proving Ground Command 
at Eglin Field, Florida, the Fourth Air Force at 
Muroc Lake, California, and at Tonopah, Nevada. 

The early work at Aberdeen was concerned chiefly 
with the development of nose cameras for recording 
the performance of the bombs, 11 ' 12 with early roll- 
stabilization experiments, and with preliminary 
tests with 100-lb bombs to establish qualitatively 
that reasonable deflections of the trajectories could 
be obtained from small control surfaces. 

The work at Eglin started in December 1942 with 
a series of bombs controlled by clock-driven cams. 
The trajectories of these bombs had been computed 


on the Rockefeller differential analyzer using data 
from the Wright Brothers wind tunnel at MIT. The 
work was plagued by mechanical failures which 
masked the significant data. Out of 10 drops 7 failed 
to stabilize in roll. From the remaining 3, however, 
important facts were learned. As shown in Figure 13 
the radius of curvature was less than expected and, 
as described in Section 2.5, the lift and maneuver¬ 
ability were therefore greater. Two of the three 
bombs which were successfully roll-stabilized were 
deflected along one coordinate only. 



Figure 13. Comparison of calculated and observed 
trajectory of cam-controlled bomb. 

The work was continued in February of 1943 with 
twelve more bombs, two of which were of the MIT 
cylindrical design referred to in Section 2.4.2 and 
shown in Figure 6. These bombs were all radio-con¬ 
trolled and while the average of success was not much 
greater than in the December tests, that of one drop, 
Figure 14, was great enough to stimulate consider¬ 
able interest in Azon. 

Characteristically, the Division questioned whether 
the single successful drop of Figure 14 was not an 
accident—the final accuracy of 20 ft rather than the 
plotted trajectory having been presented. Accord¬ 
ingly, succeeding drops were made in pairs simulta¬ 
neously released: one standard bomb uncontrolled, 
and one Azon or Razon to which control was applied 
(Figure 15). This technique gave an approximate 
measure of the error corrected by radio control with- 





44 


AZON AND RAZON 


out waiting for the analysis of motion-picture camera 
records. 

A word should be said about the presentation of 
trajectories. No satisfactory frame of reference has 
been developed for defining the trajectory of a guided 
bomb. The logical system would be Cartesian, in 
which one plane would be the ground, the second 
would be a vertical plane containing the bombing- 
run, and the third would be a vertical plane normal 
to the second and containing the target. Instrumen¬ 
tation to yield such projections of the trajectories is 













6500 

6000 H 

*00 S 
?ooo | 

4500 J- 
< 

4000 j 
u 

DC 

3500 2 

0 

3000 u 

ul 

2^00 2 
~p> 2000 | 

IJOO 

I00O 

500 




IMPACT 







ROAD TARGET 

12 FEET WIDE 










< 













> 












\ 

> 









( 








































-< 













































1 

30 60 40 20 0 20 40 60 60 

DEFLECTION FROM 

CENTER OF ROAD - FEET 


Figure 14. First successful Azon drop. 


difficult to assemble and to coordinate. Three photo¬ 
theodolites would probably be a good solution, but 
the time coordination of three ground stations and an 
airplane is extremely difficult. In this work the tra¬ 
jectories are presented as ground projections from 
the aircraft, unaccelerated flight after release being- 
assumed. Such projections are simple to instrument 
but their true significance is sometimes elusive. A 
more complete discussion of the measurement and 
analysis of trajectories is given in Chapter 8. 

The development test program of Razon was 
largely carried out at Tonopah, with final ballistic 
experiments to establish the constants for Jag per¬ 
formed at Wendover. 


210 PRODUCTION ENGINEERING OF 

AZON AND RAZON 

However competent a development program may 
be, it is inevitable that changes in the design are 
required to suit the techniques of mass production. 
Contracts OEMsr 1081, 1258, and 1415 were made 
with the Union Switch and Signal Company to im¬ 
plement the development of the Gulf Research and 
Development Company and the research of MIT for 
combat use. The activities under these contracts are 
described in Chapter 11. 

211 AZON IN COMBAT 

A squadron of B-17’s took the first Azons into 
combat in February 1944. This group went to the 
Mediterranean Theater of Operations to join the Fif¬ 
teenth Air Force. They were accompanied by a Tech¬ 
nical Aide from the Division. This group had the usual 
difficulties which occur when new equipment is intro¬ 
duced into combat. The radio receivers (see Chapter 
6) particularly gave trouble. Nevertheless, successful 
missions were flown against the Ancona-Rimini rail¬ 
way bridges, locks and bridges at the Iron Gate on 
the Danube, and other targets. One spectacular suc¬ 
cess was the demolition of the Avisio Viaduct south 
of the Brenner Pass. This viaduct was a key route for 
personnel and supplies supporting the German de¬ 
fenses in Italy. 

As a result of this success a heavy procurement 
program of Azons was instituted, and a second squad¬ 
ron, B-24’s, was prepared for the Tenth Air Force in 
the China-Burma-India Theater. At the suggestion of 
the Division, this squadron was diverted in April 
1944 to the European Theater of Operations (ETO) to 
join the Eighth Air Force. They were successful in 
demolishing bridge targets in Normandy both on the 
Seine and Loire. Figure 16 is a strike photograph of a 
bridge at Tours successfully attacked on June 6, 
1944. The number of craters adjacent to the bridge is 
testimony to the preceding unsuccessful attempts 
against this target. 

Production orders were again increased and a ma¬ 
jor training program established. This program was 
completely cancelled in July 1944. 

In November 1944 the Division was requested to 
send a technical representative to Burma to assist an 
Azon squadron stationed there with the Tenth Air 
Force. This squadron was sent out to replace the one 
which had been diverted to the Eighth Air Force in the 




































AZON IN COMBAT 


45 



ETO in the preceding April. They had gone without 
the knowledge of the Air Communications Officer, 
within whose responsibility guided missiles rested. 
This squadron succeeded in cutting all Japanese lines 
of communication in Burma during December 1944 
and January 1945. (See frontispiece.) 

Expedited action was requested for additional 
Azon control equipment for half the heavy-bomber 
strength in the CBI Theater, and the production pro¬ 
gram was reinstituted. 


It is not clear why the Azon program was cancelled 
in the summer of 1944. There seems to be no question 
that the system whereby bombardment groups re¬ 
ceived credit for tons of bombs dropped rather than 
for targets destroyed seriously biased theater com¬ 
manders in favor of mass salvos as against aimed 
single drops. This system, together with the basis for 
award of decorations, has been harshly criticized by 
the Air Forces Evaluation Board of the Pacific Ocean 
Area. 13 A study to learn why the Azon program was 























46 


AZON AND RAZON 



Figure 16. Bridge at Tours demolished by Azon 














RAZONS AGAINST ARMORED TARGETS 


47 


terminated and to make recommendations as to the 
future policy concerning this missile was made by the 
Guided-Missile Subcommittee of the Joint Commit¬ 
tee on New Weapons and Equipment of the U. S. 
Joint Chiefs of Staff. The results of this study have 
not been made available to the Division. 

212 TARZON 

As the war closed the Division had under way the 
development of Razon-like controls for the 12,000-lb 
British Tallboy, a deep-penetration bomb. Work had 
not progressed to a point which would justify report¬ 
ing it here. It is covered in the final report 14 of the 
contractor who is continuing to serve in a consulting 
capacity on the project, which has been transferred 
to the Air Forces. 

2 is RAZONS AGAINST ARMORED 
TARGETS 

The increased deck armor of capital ships makes 
them nearly invulnerable against attack with 1,000- 
lb GP bombs. The large specific gravity of the armor¬ 
piercing or semi-armor-piercing bomb, however, 
makes it an unpromising missile for control in the 


Razon manner. Furthermore, the pitching and yaw¬ 
ing required for control would reduce the striking 
velocity which is the missile’s strength; also, there 
is a high probability that such a bomb would strike 
obliquely, losing its armor-piercing power. 

The Division requested Division 8, Explosives, to 
study the possibility of using the Munroe effect of a 
shaped charge in a 1,000-lb GP bomb casing. Their 
study and scale-model tests indicated that such a 
bomb could defeat 11 in. of armor without loss of 
blast effect. Full-scale firing tests were made on April 
1, 1945, at Dahlgren Naval Ordnance Station. The 
bomb was statically fired against the target, which 
consisted of one 11-in. and one 4-in. armor plate and 
three %-in. mild-steel plates. Each plate was sepa¬ 
rated from the next by an 8-ft air space. Some un¬ 
fuzed 100-lb bombs were stacked between the second 
and third plates of mild steel. The jet from the bomb 
penetrated all the plates of the target and detonated 
some of the 100-lb bombs. 

Consideration was given to the design of a com¬ 
posite bomb having a shaped charge with a follow- 
through explosive charge. The problem of preventing 
detonation of the follow-through when the shaped 
charge is fired will require considerable effort if a 
solution is to be found. 



Chapter 3 


FELIX AND DOVE EYE 


INTRODUCTION 

T he idea of a missile so endowed with properties 
that it would of itself seek out its target on ac¬ 
count of some character inherent in the target has 
appealed to romancers for centuries. The magic bul¬ 
let of the Freischiitz legend, for example, was so 
charmed that it would inevitably strike the heart of 
a traitor irrespective of any lack of marksmanship 
on the part of the rifleman. The radar-homing mis¬ 
siles of Chapter 1 and the photoelectric-homing con¬ 
trols described in later chapters go only part way in 
meeting this ancient need. In each case the homing 
action of the missile is determined by reflected energy 
derived from electromagnetic illumination of the tar¬ 
get—in the centimeter wavelengths in the case of 
radar, in the visible range in the case of the various 
photoelectric devices investigated by the Division. 

Between these two wavelength bands lies the group 
of radiations generally known as heat. The source of 
such radiations is inherent in the temperature and 
surface quality of every object. If, then, a method 
could be developed to identify a military target by 
its heat radiation and then to cause a bomb to home 
on these radiations, a weapon possibly less subject to 
jamming or camouflage countermeasures might re¬ 
sult. This was the objective of the Felix program. 

Considerable confusion has been prevalent in dis¬ 
cussions of infrared and heat-actuated devices. To 
aid in reduction of this confusion the Division issued 
a memorandum to contractors and liaison officers 
setting forth the properties of electromagnetic radia¬ 
tions from the upper limit of visibility, say 0.80 /x to 
approximately 20 /x. This memorandum is included 
as Appendix C of this report, and its careful study 
by those interested in heat-homing missiles is urged. 

The masking of visibility by fog is well known. 
The absorption of radiation by invisible water vapor 
is less generally appreciated, although the spectacu¬ 
larly greater visibility of arid areas compared with 
coastal regions is relatively commonplace. In addi¬ 
tion the advent of infrared-sensitive emulsions for 
photographic work has, for the lay mind, confused 
the situation by investing in any “invisible light” the 
property of prodigious penetrating power. 

This simply is not the case. The true situation can 


hardly be stated concisely, and in addition to Appen¬ 
dix C of this report the reader is referred to the work 
under Division 16, NDRC, at Harvard University. 1 
In general it can be said that so far as visible ob¬ 
structions to radiation such as mist, fog, and clouds 
are concerned, there is no wavelength radiated by a 
military target that has appreciably greater penetrat¬ 
ing power than the visible band between, roughly, 
0.40 ju and 0.70 /x- For haze, smoke, and fine dust 
there is some gain in penetration from the use of 
wavelengths which are long compared with the par¬ 
ticle diameter. 

In addition to the absorption by the agencies just 
cited there is further attenuation of radiation by at¬ 
mospheric water present as pure vapor. This absorp¬ 
tion is a function of wavelength and is at a high 
value for radiations adjacent to the visible portion 
of the spectrum, 0.7 /x to 7 /x- It is in this near infra¬ 
red region, closely adjacent to the visible, that the 
common infrared devices—photographic emulsions, 
photoelectric cells, and photochemical devices—op¬ 
erate. Indeed their generic names suggest their char¬ 
acter as being light-sensitive rather than heat-sensi¬ 
tive systems. At very long infrared wavelengths the 
radiation can be detected and measured 2 by micro- 
wave techniques. 

The Stefan-Boltzmann law (see Figure 1 of Ap¬ 
pendix C) teaches that but little energy is radiated 
from targets at these short wavelengths. At longer 
wavelengths (8.5 m to 15 /x), however, the situation 
is widely different. At these wavelengths, objects at 
ordinary temperatures radiate at their maximum 
energy level; furthermore there is at precisely these 
wavelengths a “window” (reduction) in the absorp¬ 
tion characteristic of water vapor. Figure 1 was ob¬ 
tained simply by multiplying the black-body radia¬ 
tion by the transmission through water vapor given 
in Figure 2 of Appendix C. 

If a bomb is to be controlled by heat radiation 
from its target, it is in the region of the water-vapor 
window that the measurements of heat must be 
made. Other wavelengths will not be received in suf¬ 
ficient quantity both because of their paucity at the 
source and because of water-vapor absorption. The 
techniques of measurement in this field have been 
well explored by astronomers; they consist chiefly of 


GENERAL 


49 


the use of the bolometer, gas thermometers similar 
to the Hayes cell, and the thermopile. 

This introductory point has been stressed because 
confusion still exists. Even after the full discussions 
and interchange of reports between groups interested 
in this problem, one still hears the report that, “The 
Germans had an infrared detector ten times as sensi¬ 
tive as ours.” This is probably not true; even if it is 



4 6 8 10 12 14 16 

WAVELENGTH IN MICRONS 

Figure 1. Energy received from radiating sources 
through a water-vapor curtain. 

true, it is not particularly significant as regards 
guided missiles. In every case where the attention 
of the Division has been drawn to such detectors, 
they did indeed have a high sensitivity but in the 
near infrared region. They have been photosensitive 
rather than thermosensitive systems. A similar and 
equally false comparison could be made between the 
sensitivity of bolometers and that of the microwave 
techniques of Dickie. 2 The sensitivity of his method 
(10 -16 watt) is many orders higher than that of the 
bolometers used in Felix. It is of no value to have a 


sensitivity of detection a hundredfold greater if the 
energy available at the frequency where the sensi¬ 
tivity is high is a thousandfold less. 

32 GENERAL 34 

Figure 2 is a photograph of the final version of the 
Felix bomb as it was produced for combat use. The 
main structural member was a standard M-44 1,000- 
lb GP bomb. Attached to the front by the nose-fuze 
thread was a false nose containing a scanning system 
to detect thermal targets and to transform the heat 
signal to an amplified electric signal. A false tail, 
mounted by means of the fin-lock thread, housed 
servomotors which operated elevators and rudders 
in response to the electric signals developed in the 
nose. In addition, the tail contained a twin gyro unit 
—one free and one rate gyro—which controlled aile- 



Figure 2. Felix bomb. 

rons to maintain the position of the missile fixed with 
respect to the roll axis. 

In the initial instant after the bomb clears the 
bomb bay, its flight is horizontal, headed toward the 
horizon. Not until some seconds have elapsed will it 
nose over sufficiently for the roll axis of the bomb to 
point toward the target (Figure 3). The mechanism, 
therefore, includes a time switch which keeps the 
elevators and rudders locked in neutral until the 
bomb has fallen to a point where the axis of scan 
intersects the ground in the vicinity of the target. 
For a drop from 15,000 ft this will be at about 
10,000 ft. 

Transverse lift is supplied by the body of the mis¬ 
sile itself in the same manner as Azon and Razon 
(see Chapter 2). Such a body of revolution is not an 
ideal airfoil and considerable angle of attack has to 
be applied in order to develop sufficient transverse 
acceleration to correct errors of aim or to follow the 
maneuvers of an evasive target. The scanning system 
was therefore mounted on gimbals and connected by 


























50 


FELIX AND DOVE EYE 


cables with the rudders and elevators so as to align 
the axis of scan more nearly with the tangent to the 
flight path (Figure 4). This coupling also served as a 
stabilizing feedback. 

The scanning system is discussed in detail in a 
subsequent section of this chapter. In general, it 
consisted of an optical bolometric assembly Avhich 
scanned the terrain in a field of 10-degree radius ap¬ 
proximately centered on the point where the bomb 
would fall with no further control added. The re¬ 
ceived heat measured by the bolometer while scan- 


25,000—" 

FT 


N. 


N 


\ 

\ 


20,000 

FT 



' \ 


15,000- 


FT 


Ns 



Figure 3. Free-fall portion and homing portion of 
Felix trajectory. 


ning opposite half-fields was compared, and the mis¬ 
sile was directed toward the half-field radiating the 
greater heat. Two channels provided one pair of half¬ 
fields in the range sense and another pair in the 
azimuth sense. 

The scanning system contained no intentional 
dead band where there was no signal for correction 
in any direction. The servomotors, therefore, were 
continuously energized for full speed in one direction 
or the other. Except for the very small interval re¬ 
quired for reversal upon instantaneous reversal of 
armature current, they operated at constant speed 
to give “up” or “down” elevator and “right” or 
“left” rudder. Such a system is inherently oscillating. 



DESICCATOR- 
GIMBALS- 
WIN DOW- 

BREATHING. 

DIAPHRAGM 


SPRING 


PROTECTION 
"CAPS 


Figure 4. Assembly of Felix. 



















































































































SYSTEM DYNAMICS 


51 


The back coupling from the rudders and elevators to 
the gimbal-mounted scanning head tended to reduce 
the amplitude of the oscillations. 

33 # SYSTEM DYNAMICS 

331 The Equations of Motion 

The motion of Felix under the control of its hom¬ 
ing device consists of two interdependent portions: 
motion of the center of gravity under the acceleration 
of gravity and the aerodynamic lift, and rotation 
about the yaw and pitch axes under the action of 
aerodynamic moments. Rotation about the roll axis 
is prevented by the gyro-aileron system as in Razon. 
In Figure 5, 

X is the horizontal range of the bomb from the re¬ 
lease point; 

H is the instantaneous altitude of the bomb; 

V is the instantaneous direction of the velocity 
vector whose value is v; 

B is the axis of the bomb in roll; 
a is the angle between the velocity vector and the 
bomb axis, the angle of attack; 

8 is the elevator displacement with respect to the 
bomb axis; 

s is the instantaneous direction of the axis of scan; 
< i > is the angle between the axis of scan and the 
bomb axis; 

H is the measured error, the angle between the 
axis of scan and the bomb-target line; 

\f/ is the true error in heading, the angle between 
the instantaneous velocity and the bomb-target 
line; 

7 is the elevation from vertical of the bomb-target 
line. 

A lift perpendicular to the velocity is developed by 
the angle of attack set up by the elevator displace¬ 
ment. In the steady state this lift is 

L = hpV 2 C L A (1) 

where p is the air density, Cl is the lift coefficient, 
and A is the area over which the lift is developed. 
The relationship between C l and a is linear for small 
values of a and for airfoils of conventional shape. 
Although for bombs of this type the departure from 
linearity is appreciable, it was neglected in this case. 
A drag in line with the instantaneous velocity is: 

D = \ P V*C D A (2) 

The relationship between the drag coefficient Cd and 


a is quadratic. The position of the center of gravity 
is determined, then, by: 

X = ~L cos (7 + t) - ~D sin (7 + i£) (3) 

and . ! 1 

~H = g - —L sin (7 + 1) ~ ~D cos (7 + (4) 

where m is the mass of the bomb. 

RELEASE 

POINT 



Figure 5. Control of Felix in range. 

The angle of attack which determines the lift and 
drag is derived from the rotation of the structure 
about its center of gravity. The rotation is reponsive 
to a pitching moment: 

M = \ p V*CmA (5) 

where the moment coefficient Cm is a function of a 
and 8. An approximate fit of the steady-state wind- 
tunnel data is given by: 

Cm = kiot + k*o? — 8 (6) 

The rotation is opposed by an aerodynamic damping 

R = k 3P V (7) 

The rotary motion then becomes 

I a + Ra — M = 0 ( 8 ) 

The simultaneous solution of equations (3), (4), and 
( 8 ) defines the motion of the missile. 

The solution is complicated not only by the non¬ 
linearity of the relationship between the forces and 
moments and the corresponding displacements but 
more particularly by the relationship between 8 and 
the observed error angle H. The elevator motion 5 












52 


FELIX AND DOVE EYE 


is discontinuous with time. Between the travel lim- 

its, ±5 mas 

8 = /c 4 

if H > 0 

(9) 

8 = — k 4 

if H < 0 

(10) 

Furthermore 



R = a 

- ^ + 7 

(H) 


where H by virtue of the coupling from the servo 
link is partly proportional to 8 and partly a time 
function developed by the bias device. 


3 3 2 On-Off Control 

An on-off control is extreme^ difficult to analyze. 
The British did considerable work in this field, ref¬ 
erences to which have been made in various Army 
reports. In addition the Germans 6 ’ 7 ’ 8 ’ 9 ’ 10 ’ 11 worked 
exhaustively on this subject. The differential ana¬ 
lyzer at MIT is inadequate to cope with the prob¬ 
lem in its entirety. Its use was invoked for simpli¬ 
fied portions, but in the main the resolutions were 
obtained by extremely laborious point-by-point in¬ 
tegration and by means of a simulator. 

The simulator is described in Chapter 10. It pro¬ 
duced solutions of the equations of single-axis rotary 
motion, including the nonlinearity in the relation¬ 
ships. In the point-by-point integration the complete 
analysis was made. In each method, however, the as¬ 
sumption was made that steady-state data taken in 



w 

w 

AXIS OF\\ 

SCAN- 

w 

\\ 

\\ 

\ 

\ 

POINT OF IMPACT WITH 
NO ADDITIONAL CONTROL 

Figure 6. Parabolic thinking. 


the wind tunnel is applicable throughout the tran¬ 
sient regime. 

While no thorough quantitative analysis of the 
dynamics of the Felix system is therefore practicable 
in this report, a qualitative discussion will be helpful 
in explaining its operation. An on-off control system 
such as this one keeps the rudders and elevators in 
continuous oscillation. If the scanning system were 
rigidly mounted with axis collinear with the roll axis 
of the bomb, the whole structure would oscillate, with 
the roll axis swinging about a line continuously 
pointed at the target. This would be an extremely 
inefficient control, since a considerable angle of at¬ 
tack is required to develop the necessary correcting 
lift, and with the oscillations of the structure the 
average angle of attack would be continuously defi¬ 
cient. Back coupling causes the rudders and elevators 
to oscillate about a mean position which causes the 
scanning system to look continuously at the target. 
Thus if a correction is called for, the control surfaces 
will oscillate about a mean position which gives lift 
in the direction to correct the error. A quasi-propor¬ 
tional control is thus effected. If the back coupling 
could be made through a linkage which matched the 
dynamic a-to-5 relationship, i.e., so that the axis of 
scan was continuously tangent to the flight path— 
then the measured error angle would become the 
same as the true error angle, 

H = i (12) 

and a truly proportional control would ensue, where 
8 is the mean control-surface displacement: 

8 = K+ (13) 

Without knowledge of the transient response of 
airframes, such a design is impossible. 

Such a control system, however, would still be 
wasteful of controllability, especially in the range 
sense. As the bomb falls in its normal parabolic path, 
the first glimpse which the scanning system will have 
of the target would indicate a gross overshoot, even 
if the bomb were perfectly aimed or were aimed 
short. In Felix a bias device was added which super¬ 
posed on the back coupling a range deflection which 
caused the scanning axis—with the elevators in neu¬ 
tral—to look along a chord of the parabola at the 
point on the terrain where the missile would fall with 
no further control added (Figure 6). This system of 
biasing, known as parabolic thinking , was deemed by 
Gulf 12 to be unnecessary. The deflection between the 
axis of the scanning system and the axis of the bomb 



SCANNING SYSTEM 


53 


(</>*) which the bias device inserts is a function of 
altitude and the range component of airspeed. Since 
the latter is subject to considerable variation be¬ 
tween the points of release and impact, the conclu¬ 
sion of the Gulf investigation was that sufficient ac¬ 
curacy could not be obtained in its estimate to justify 
correcting for it. The success with the photoelectric 
high-angle target-seeking bomb (see Chapter 9) sup¬ 
ports their conclusion. Further study is needed. 

34 SCANNING SYSTEM 

3-4,1 Bolometer 

The requirements for the thermosensitive element 
in the Felix system are high sensitivity, freedom from 
microphonics, and low time constant. Thermopiles 
were tested but found to be inferior to bolometers, 
although General Motors in an investigation made 
for the AAF is said to have had some success with 
far infrared detectors of this type. Bolometers of 
many types were tested, including the nonmetallic 
thermistor of Bell Telephone Laboratories. None was 
more satisfactory than the pure-metal type, nickel or 
gold. The former was the type finally used in the pro¬ 
duction design. Work was done with gold bolometers 
consisting of a single strip of evaporated gold sup¬ 
ported on a thin nitrocellulose diaphragm. 13 Although 
these bolometers worked well in drop tests, they 
seemed less suited to manufacture than the nickel 
strip bolometers developed at MIT. 

The bolometer element proper consisted of four 
nickel strips 0.26x0.045x0.00001 in. The nickel 
strips were manufactured 14 ’ 15 by electroplating nickel 
from a hot (90 C), concentrated solution of nickel 
ammonium sulphate onto an aluminum cathode. 
Thickness was controlled by adjustment of the time 
of current flow. While supported on the cathode the 
nickel film was cut to the correct width (0.045 in.), 
after which the aluminum was removed by dissolv¬ 
ing it in a 1 per cent solution of sodium hydroxide. 
The plated cathode, scarred by the cuts in the nickel, 
was immersed in a shallow bath of the solvent. The 
nickel strips quickly floated to the surface, borne by 
the bubbles formed as the aluminum dissolved. The 
strips were then washed in distilled water and dried 
flat on paper. 

The strips were supported in the bolometer (Fig¬ 
ure 7) by springs of phosphor bronze, designed so as 
to produce a natural period in the nickel strips well 
above the scanning frequency (32 c). This required a 


pull of 2 to 3 grams—equivalent to a loading of 
about 12,000 psi in the nickel. Some damping was 
provided by making one set of the phosphor bronze 
springs on one end softer than the other. The nickel 
strips were hard-soldered to the springs. 

The springs were supported on mounting and lead- 
in wires in a glass-steel press to which a cap of pure 
silver was soldered. The top of the cap had a dia¬ 
phragm opening 0.199 in. in diameter, and the whole 
was covered with a drawn spherical-segment window 
of pure silver chloride fused to it. To increase the 
speed of response of the bolometers, the assembly was 



filled with dry hydrogen at approximately 3 mm 
pressure. 

The sensitivity was increased by covering the front 
surface of the nickel strips with a coating of gold 
black formed by evaporating pure gold onto the 
nickel in a low-pressure atmosphere of hydrogen. The 
technique required nice adjustment since the coating 
to be effective must have a high absorption in the 
8.5- to 15-jz region combined with a low thermal mass. 
A good coating increases the thermal sensitivity of 
the bolometer by a factor greater than 5. Too heavy 
a coating can increase the thermal mass and thus 











































54 


FELIX AND DOVE EYE 


reduce its effective sensitivity. Full instructions are 
given in Appendix B of the final Felix report. 14 

3 4 2 Optics 


The optical system (Figure 8) consisted of a rotat¬ 
ing parabolic mirror of approximately 1-in. focal 
length and aperture of 2J^-in. diameter. The axis of 



the parabola was inclined 5 degrees to the axis of 
rotation, the axis of scan. The two axes, that of 
rotation and that of the parabola, intersected at the 
focal plane of the mirror. The bolometer was mounted 
with its center at the intersection of the axes. The 
diaphragm in the bolometer assembly (see Section 
3.4.1) limits the field which can be projected onto 
the bolometer strips to a circle 10 degrees in diam¬ 
eter, 5 degrees either side of the parabola axis. 




Figure 9. Scanning pattern (top); relative sensitivity 
(bottom). 

The combination of the diaphragm and the inter¬ 
section of the optical axis with the axis of scan as¬ 
sured that no rays more than 5 degrees off the para¬ 
bolic axis were used. As the mirror rotated, rays par¬ 
allel to the rotational axis focused at a point which 
traveled around the edge of the diaphragm aperture. 
Rays 5 degrees off axis crossed at the center of the 
bolometer; rays 10 degrees off axis just intersected 
the edge at a point diametrically opposite to that of 
the ray incident along the axis of scan. 

A simpler way to look at such a system is to con¬ 
sider it in reverse, with the bolometer as a radiator 





































SCANNING SYSTEM 


55 



rather than as a receptor. Thus an image of the dia¬ 
phragm is projected on the terrain by the mirror. 
As the mirror rotates the image nutates—i.e., trans¬ 
lates in a circle—about the axis of scan. The distinc¬ 
tion between nutation and rotation is important, for 
if the image of the bolometer diaphragm could be 
made to rotate (which could be accomplished, for 
example, by a rotating bolometer) the sensitivity of 
the system to targets near the edge of the field could 
be increased by using a cardioid-shaped aperture. 

Figure 9 shows the sensitivity of the scanning sys¬ 
tem as a function of measured error angle. The de¬ 
parture of this characteristic from the ideal square 
curve (shown dotted) is the result of three causes. 
The first cause is due to the nutational character of 
the scanning. The scanning pattern shows the tracks 
of targets at 1 degree and 9 degrees off axis. It shows 
that the target near the center is on the bolometer 
for 49 per cent of a scanning cycle; the more remote 
target is on the bolometer for only 15 per cent of the 
cycle. 

The second cadse is inherent in the optics of a fast 
parabola, which produces a good image on the axis 
but a confused one from an object 5 degrees off axis. 
Schmidt optics could correct this, but mounting a 
Schmidt corrector plate in such a scanning system 
presents a serious mechanical problem. 

The third cause is the essentially flat character of 


a hole. This makes the diaphragm aperture, which 
looks like a circle when viewed from the center of the 
mirror, appear as an ellipse when viewed from the 
mirror’s edge. With the aperture ratio used (/ 0.36) 
the apparent axes of the ellipse were 10 degrees and 
3.5 degrees. As a result, the effective speed of the 
parabola is much reduced for off-axis rays. The only 
cure for this would be reduction of the parabola 
aperture, with consequent decrease in the overall 
sensitivity. Further, the large parabola served an¬ 
other purpose as a shield for thermal radiations from 
the scanner case. 

Figure 10 shows a cross section of the complete 
scanning unit. The mirror surface and the bolometer 
have to be protected against any condensation. This 
was accomplished by sealing the mirror and bolom¬ 
eter in a dry air space by means of a conical window 
of 3J/£-in. base diameter and having an apex angle of 
136 degrees. The conical shape and apex angle were 
so chosen as to prevent any heat from the warmed 
bolometer or its case from being reflected onto the 
parabolic mirror and back onto the bolometer strips. 

To prevent fogging of the external window surface, 
the window material, silver chloride, was made Yi. 
mm thick, giving it a low thermal mass. Silver chlor¬ 
ide is highly conductive (80 per cent transmission) 
to wavelengths between 8.5 and 13 \i. It is, however, 
extremely sensitive to actinic light. In the strong 



















































































56 


FELIX AND DOVE EYE 


ultraviolet illumination found at altitude it becomes 
nearly opaque in the visible and partially opaque to 
infrared in less than an hour. This difficulty was sur¬ 
mounted by coating the windows with silver sulfide in 
a layer sufficient to reduce the transmission in the 
0.4- to 2.0-M band to about 1 per cent while preserv¬ 
ing a transmission in the 8.5- to 13 -m band of 70 per 
cent. Uncoated windows have a transmission in this 
band of about 80 per cent. 

A second difficulty with silver chloride is its ex¬ 
treme corrosive tendency when in contact with base 
metals. This was curbed by the use of a nonmetallic 
gasket between the window and the scanner case. 



Figure 11. Scanning system in place. 


The large window was capable of standing a pres¬ 
sure differential of approximately 4 psi, which is less 
than the sum of the dynamic pressure and the pres¬ 
sure change from high altitude to sea level. An equal¬ 
izing system was constructed by sealing the bulkhead 
which closes the back of the scanner compartment, 
and the annular space between the front of scanner 
and the nose fairing was closed with a flexible, porous 
gasket of nylon. Thus the whole cavity housing the 
scanner (Figure 11) was kept close to the same pres¬ 
sure that existed on the nose of the bomb. The inside 
of the scanner—the chamber behind the silver chlor¬ 
ide window—was connected with this cavity through 
a large tube containing silica gel as a desiccant. 


The remaining elements of the scanning head con¬ 
sisted of the driving motor and the commutator. The 
motor was of a special design with close tolerances to 
permit direct mounting of the parabolic mirror on 
the motor shaft with preservation of optical align¬ 
ment. To reduce magnetic pickup, the shaft was 
made of nonmagnetic material (bronze). To prevent 
change in phase shift as the signal was amplified, the 
motor was regulated at 32 rps by a centrifugal gov¬ 
ernor. A double-arm design was adopted to make the 
governor insensitive to gravity. With the usual single¬ 
arm design, the governor tends to cut in and out at 
scanning frequency whenever the shaft is not verti¬ 
cal. The change in field surrounding the motor caused 
considerable pickup in the amplifier, which was tuned 
for this frequency until the double-arm design was 
adopted. 

A commutator of the simple cam and rocker-arm 
type, mounted on the back plate of the motor, indi¬ 
cated the phase of the bolometer signal. Each of its 
four contacts—up-down and right-left—was set to 
close through 165 degrees of rotation; a small adjust¬ 
ment was provided to accommodate variation in 
phase shift through the amplifier. 

35 ELECTRONICS AND SERVO LINK 
3,5,1 Principles of Operation 

The heat signal from the target is intermittently 
focused on the bolometer as described in the preced¬ 
ing section. When the heat strikes the bolometer, its 
temperature rises rapidly; as the heat image of the 
target leaves the bolometer, its temperature falls ex¬ 
ponentially until the image falls on it in the next 
scanning cycle. This fluctuating temperature is ac¬ 
companied by a fluctuating resistance so that the 
bridge circuit of Figure 12 is cyclically thrown off 
balance. The transformer, the primary of which 
forms two arms of the bridge, serves as an impedance 




Figure 12. Bolometer circuit. 






















ELECTRONICS AND SERVO LINK 


57 


match—about 10 ohms—between the bolometer and 
the high-impedance grid input of the first stage. 
Figure 13 represents the temperature change in the 
bolometer. In addition to providing an impedance 
match, the transformer, by eliminating the d-c com¬ 
ponent and blocking the higher harmonics, converts 
the signal to the approximately sinusoidal form shown 
opposite the input circuit element in the block dia¬ 
gram of Figure 14. The preamplifier lifts the signal 
to a usable level and further removes harmonics. 

The preamplifier is followed by a twin-triode stage 
which produces phase inversion, resulting in a two- 
channel output whose voltages are equal but in phase 
opposition. The bias on this stage is such as to pro¬ 
duce saturation (clipping) for strong signals. A sec¬ 
ond clipping stage squares up the signal from weak 
inputs and maintains the amplitude independent of 
the strength of the input. The output of the phase 



TIME -► 

Figure 13. Temperature variation in bolometer. 

inverter stage is a pair of chopped waves; the output 
of the clipper is a pair of square waves. 

The signals from the two channels are now ana¬ 
lyzed and resynthesized by the commutator. Consider 
the situation represented by the block diagram of 
Figure 14 where the target is on course in the azi¬ 
muth sense but low. At the beginning of a scanning 
cycle the output of channel A is positive, that of 
channel B is negative. During the first quarter-cycle 
the right-left channel receives its voltage through the 
commutator from channel B (negative). During the 
half-cycle it receives its voltage from channel A (posi¬ 
tive in quarter-cycle 2, negative in quarter-cycle 3). 
During the final quarter-cycle it receives its voltage 
from channel B, which has now swung positive. The 
output to the rudder channel is therefore a balanced 
square wave of a frequency twice the scanning fre¬ 
quency. 


The situation as regards the elevator channel is 
quite different. At the first instant after the begin¬ 
ning of the scanning cycle it receives its voltage from 
channel A (positive). This condition continues for 
the complete half-cycle, at the end of which the volt¬ 
age supply is switched to channel B, which has now 
bepome positive in turn. The commutator output to 
the elevator, then, consists of a substantially constant 
positive voltage. 






^inri rL 

Tjuinr 


\* \ a \ /? \ & : 


tJUuMIUI 




Figure 14. Block diagram of Felix electronic servo 
system. 


The commutator outputs were fed into resistance- 
capacitance circuits which integrated the voltage over 
several scanning cycles. Thus, for the condition 
shown in Figure 14 the input to the elevator-control 
circuit slowly climbs. When the output of the inte¬ 
grator circuit, amplified by the control tube, reaches 
the pickup voltage for the control relay, the relay 
operates. The elevator servomotor reverses, depress¬ 
ing the elevators until the feedback connection to the 








































58 


FELIX AND DOVE EYE 


scanner moves the target into the top half of the field 
of view, when the whole cycle is reversed. 

The input to the rudder-control circuit oscillates 
at a low amplitude about zero. The relay therefore 
remains on the “right” contacts. It should be noted 
that the diagram is somewhat incomplete, for after 
a few scanning cycles the rudder servomotor would 
cause the scanner to deflect, bringing the target into 
the “left” half of the field of view. It is this feedback 
operation which gives the quasi-proportional con¬ 
trol discussed in Section 3.2. 

3 5 2 Transformer and Input Circuit 

The transformer was of special design, having two 
59-turn sections in the primary and 6,500 turns in 
the secondary. It was wound on a Mu-metal core and 
provided a net gain of 77 at 32 c. Three cases of Mu¬ 
metal and one of copper produced a shielding of 
90 db. 

The overall size was V/% in. in diameter by 2% in. 
long. The final version of the missile provided more 


$ © 


0 *9 



Figure 15. Bolometer unit and its components. 

space than this for the transformer. A redesign might 
have improved both the operating characteristics and 
its ease of production. Transformer design talent was, 
however, one of the most critical procurement prob¬ 
lems which faced the Division, and the revision of a 
working design was not undertaken. 


The two halves of the transformer primary formed 
two arms of a bridge circuit; they were balanced to 
within 1 per cent as to resistance and impedance at 
32 c. The bolometer comprised a third arm, which was 
balanced by a special resistor wound with Hytemco 
wire, an alloy of high resistivity having about the 
same temperature coefficient of resistance as the 
nickel strips. Close balance was essential to eliminate 
noise and to prevent saturation of the transformer 
core. 

Power for the bridge was supplied from a special 
6-v lead-acid battery connected to the bridge by 
short shielded leads. This portion of the circuit op¬ 
erated at the lowest power level, approximately 1.0 
microvolt, and extreme care had to be taken to mini¬ 
mize pickup. The bridge type of circuit reduced some¬ 
what the voltage output available from the bolom¬ 
eter but it reduced the noise by a much larger factor, 
the overall result of the circuit being a gain of about 
50 in signal-to-noise ratio. Taken as a whole the in¬ 
put circuit acted as a broad-band-pass filter, peaked 
at 32 c and having half-power points at 18 and 60 c. 
The low cutoff is determined by the transformer it¬ 
self, the higher limit by a 0.005-juf condenser on the 
transformer output. 

The transformer, series resistor, and Hytemco bal¬ 
ance resistor were housed in a heavy iron shield, 
shock-mounted to guard against microphonics, which 
constituted an exceedingly acute problem. All con¬ 
necting wires had to be carefully tested against the 
production of noise from flexure. Each joint in the 
primary circuit had to be made so that no conductors 
touched except within the solder bead of the joint. 
The sensitivity of the bridge circuit to changes in 
resistance, one part in one hundred million, imposed 
these rigorous specifications. Accordingly the entire 
input circuit (Figure 15), bolometer, resistors, and 
transformer, was made up as a unit, terminating in 
a 4-prong plug for connection to the battery and 
amplifier. 

3 5 3 Preamplifier 

The preamplifier consisted of two resistance- 
coupled pentodes with fixed bias. The usual objection 
to the use of fixed bias, wide variation in tube per¬ 
formance with fluctuations in plate-voltage level, was 
avoided by stabilizing the plate and screen voltage 
supply with voltage-regulating tubes. This had the 
advantage in such a low-frequency amplifier of avoid¬ 
ing large, oil-filled, cathode-by-pass condensers. 



ELECTRONICS AND SERVO LINK 


59 



The band-pass of the preamplifier was approxi¬ 
mately 45 c, peaked at 32 c. It is not a particularly 
narrow band but it is sufficiently narrow to eliminate 
a large portion of the noise, that above approximately 
60 c. It could not have been made much narrower 
without introducing serious phase shift with varia¬ 
tion in scanning frequency. 

3 5 4 Phase Inverter and Clipper 

Phase inversion was accomplished by a twin triode 
(12SL7), with the input applied to one grid and a 
common cathode resistor. The second grid was ef¬ 
fectively grounded to voltages of 32 c through a 0.5-juf 
condenser. Thus the two plate outputs were 180 de¬ 
grees out of phase. These two plate circuits form the 
start of channels A and B of Figure 14. The second 
twin triode (12SN7) of Figure 16 clips or squares the 
signal. Both the phase inverter and the clipper op¬ 
erate with a positive 26-v bias on the grid, which 
both stabilizes the operation of the amplifier and 
prevents overloading and blocking of the circuit from 
strong signals. This is most important, since blocking 
even for a fraction of a second would seriously upset 
the operation of the integrating circuit. 

The whole scheme of operation of the electronic 
servo system is based on phase discrimination. In the 
preamplifier phase distortion was avoided by making 
the band-pass wide enough to take care of scanning- 


frequency variations. The signal level in that portion 
was so low that danger of phase shift from amplitude 
distortion existed. In the clipping section, however, 
the reverse is the case. This section had to have a 
wide band-pass to produce a square-wave output so 
that there was no danger of phase shift due to fre¬ 
quency swings. Phase shift from amplitude distortion 
was avoided only by careful design. Since the very 
essence of the clipping-circuit operation was a sym¬ 
metrical distortion of the incoming signal, a careful 
balance between the positive and negative portions 
of the square-wave output had to be maintained. 

3 5 5 Commutator and Integrator 

The action of the commutator has already been de¬ 
scribed (see Section 3.5.1). It synthesized the signals 
in channels A and B to produce the control signals 
for the elevator channel and the rudder channel. Its 
output was used to charge or discharge a 0.25-/d con¬ 
denser in each of the two control channels. The high 
side of each condenser was connected to a grid of the 
twin-triode control tube which had a 3-v negative 
bias; a potential shift of 0.5 v at the control-tube 
grid would cause the relays to operate. 

The length of time required for the relays to re¬ 
ceive an operating signal depends on the voltage to 
which the input condenser at the control-tube grid 
was last charged. It swings during operation between 





























































































60 


FELIX AND DOVE EYE 


0 and —6.0 v. The time constant of the input circuit 
is such as to require about 5 scanning cycles fully to 
charge or discharge the condenser. 


Figure 17 indicates the operation of the system 
with target motion from the down-left quadrant to 
the up-right quadrant. This figure assumes that the 



OUTPUT OF 
CHANNEL A 



OUTPUT OF 
CHANNEL B 


OUTPUT FROM 
COMMUTATOR 
TO U-D CHANNEL 



OUTPUT FROM 
COMMUTATOR 
TO R-L CHANNEL 





Figure 17. Operation of Felix electronic servomechanism with target motion from down-left quadrant to up-right 
quadrant. 






















































































































































FIELD TESTS 


61 



DIRECTION OF VERTICAL PROJECTION 
PLANE FOR TRAJECTORY 


timp im ^Fr.nwn<i- 


PROJECTION OF PATH 
ON GROUND 



. TARGET 
9 IMPACT 


ANGLE - 5.95' 


RANGE IN THOUSANDS OF FEET- 


Figure 18. Plan and profile trace of successful Felix drop at Eglin Field, January 1944. 


target has been in the down-left quadrant for a suf¬ 
ficient interval for the integrating condensers in the 
elevator channel and the rudder channel to have 
stabilized at 0 and — 6 v respectively. The target is 
assumed to cross the intersection of the axes at t = 0. 
Ten scanning cycles later the condenser charges have 
changed sufficiently to cause relay operation. 

3 5 6 Servomotors and Feedback 

The servomotors controlled by the output relays 
of the electronic servomechanism were simple 24-v 
shunt-wound motors. The fields were continuously 
energized. Their armatures were energized for clock¬ 
wise or counterclockwise rotation, depending upon 
whether the control relays happened to be resting on 
their front or back contacts. As has been already 
pointed out, this gives rise to oscillation of the con¬ 
trol surfaces about a mean displacement essentially 
proportional to the error in heading. 

The motors were geared to give a displacement rate 
of 12 degrees per second to the elevators; this speed, 
combined with the time constant of the integrating 
circuit and the feedback ratio, the ratio of 8 to <£, 


determined the frequency of oscillation (about 1.2 c) 
of the control surfaces. 

Increasing the feedback ratio and the servomotor 
speed or decreasing the time constant of the integrat¬ 
ing circuit would raise the frequency of control- 
surface oscillation. This might result in more stable 
flight at the expense of a heavier power demand by 
the servomechanisms. In any case this frequency 
must be kept reasonably remote from the natural 
frequency of the missile in pitch, as determined by 
the partial derivative of the pitching moment at 
trim and the moment of inertia. The investigators on 
this project chose to keep the period of control- 
surface oscillation well above the natural period of 
the missile. The restricted space in Felix for batteries 
and servomotors speaks strongly for this choice. 

3 6 FIELD TESTS 

3 61 Eglin Field Tests 

The initial tests with an American heat-homing 
missile were made at Eglin Field in January 1944. 
Six bombs of a preliminary design were prepared, 










































62 


FELIX AND DOVE EYE 


based on the design of the Gulf photoelectric target¬ 
seeking bomb (see Chapter 9). Flight tests had been 
made with the scanning head with the control relays 
energizing indicating lights. The results of these tests 
seemed to indicate promise of control by means of far 
infrared radiation. The results of the Gulf tests with 
the photoelectric target seeker indicated that the 
high-angle bomb had sufficient dirigibility to permit 
homing control. 



Figure 19. Impact pattern of 12 Felix drops, Septem¬ 
ber 1944. 


A target was constructed by clearing an 800-ft 
square area surrounded by scrub undergrowth after 
the first drop indicated that the regular target at 
Range 55 was an inadequate radiator. Of the six 
bombs dropped, two failed to stabilize in roll. The 
remaining four homed or tried to home on a target 
which subsequent survey showed to be a heat radia¬ 
tor. One of them made a successful homing-flight 
landing some 30 ft from the center of the target. 
The ground projection and profile of its trajectory 
are given in Figure 18. 

These results, while not spectacular, were suffi¬ 
ciently encouraging to enlist the strong support of 


the Air Forces Communications Officer, who had 
been given the added duties of serving as senior liai¬ 
son officer for the Division. Intelligence had disclosed 
targets of high priority, which later proved to be 
launching sites for the German V-l; accordingly the 
Air Forces urged utmost acceleration of the Felix 
program as a weapon against these well-camouflaged 
targets, which were known to contain diesel engines 
and air compressors. It was hoped that this ma¬ 
chinery might prove to be a sufficient source of heat 
to make Felix an effective weapon against them. 

Accordingly NDRC accepted in March 1944 the 
assignment of crashing the Felix program, although 
it was clearly understood that the targets might 
prove submarginal and that the chances of proving 
a combat design within a year were indeed remote. 

3 6 2 Tonopah Tests 

Two designs were rushed; one, having a cruciform 
empennage permitting better stowage in bombard¬ 
ment aircraft, was terminated when the experience of 
Gulf with Razon (see Chapter 2) indicated that roll 
stabilization with a cruciform tail was hardly to be 
accomplished. The second design with an octagonal 
empennage was pushed to completion. Laboratory 
prototypes were ready for drop tests by August 1944. 

The initial target consisted of aluminum foil nailed 
to panels laid out in the form of a square cross on the 
desert floor. The arms of the cross were 40 ft by 100 
ft, the total circumscribing square being 240 ft by 
240 ft. It was hoped to make the bombs home on the 
sun’s reflection from this target; however, it proved 
to be wholly unsatisfactory. While the signal from 
the sun’s reflection was very large indeed, missions 
could be flown only in the middle of the day, when 
the elevation of the sun was approximately the same 
as the bomb after falling to an altitude of 10,000 ft 
from a release point of 15,000 ft. The aircraft had to 
fly directly into the sun in order for the scanner to 
pick up a good reflection. The greatest disadvantage, 
however, was that in the middle of the day the air 
was so turbulent that the bombardier was unable to 
get a good approach run. Eight drops against this 
target proved abortive. 

A second target was constructed, consisting of 100 
rectangular steel plates, each approximately 30 sq ft 
in area. These plates were mounted so that they in¬ 
clined 30 degrees from the vertical and were heated 
to a very dull red heat with 3 oil burners (orchard 
heaters) per plate. Eleven Felix bombs were dropped 





FIELD TESTS 


63 



at this target. Seven displayed electrical malfunc¬ 
tions. Four made successful homing flights, scoring 
misses of 100 to 300 ft. 

The design was then subjected to an engineering 
analysis. Longer rudders and elevators were installed, 
producing a larger angle of attack and permitting 
the elimination of a troublesome lift shroud. The 
control relays were redesigned and housed in a her¬ 
metically sealed case. The rudder deflection speed 



was reduced about 30 per cent to a final value of 12 
degrees per second to remove the range of forced 
oscillation from the vicinity of the natural frequency 
of the bomb. The feedback coupling mechanism was 
simplified and the bias device to provide parabolic 
thinking was introduced. The integrator stage was 
added to the amplifier. 

Sixteen of these bombs were tested in September 
1944. Twelve of the sixteen homed successfully, pro¬ 
ducing the impact pattern shown in Figure 19. Com¬ 
posite plots of the azimuth and profile traces of the 
trajectories are shown in Figures 20 and 21. 


Figure 20. Azimuth traces of Tonopah drops. 















64 


FELIX AND DOVE EYE 


As a result of these reasonably successful tests the 
designs were released to the manufacturer for a pre- 
production order of 100 units. 

3 6 3 Ocala Tests 

The preproduction units from the manufacturer 
were desired by the Air Forces Board for evaluation 
at Orlando, Florida. The Division was reluctant to 
release the initial units from a new supplier for this 
purpose; accordingly a compromise was worked out 
whereby the initial phase of the test program to 
check the manufacturing design would be carried 
out by the Division’s contractor. 

It was now obvious that the design of a heat tar¬ 
get was a critical element of the testing program. 
The Division undertook to construct a target to 
simulate thermally an operational target in which 
the Twentieth Air Force had a strong interest. Figure 
22 is an aerial photograph of the target constructed 
at Ocala range near Orlando. The target consisted of 
eleven areas cleared of vegetation to represent the 
eleven buildings of a Japanese aircraft-engine fac¬ 


tory. Oil burners similar to those used at Tonopah 
were installed to augment the reradiated solar energy. 
Surveys made in November indicated adequate con¬ 
trast from the surrounding vegetation. 

The tests began on December 27, 1944. It was im¬ 
mediately apparent that a great many defects had 
crept into the manufacturing procedure and that in¬ 
spection methods adequate to detect them had not 
been developed. The hope that any of the preproduc¬ 
tion units could be made available for evaluation 
soon seemed naive. Too much credit cannot be ex¬ 
tended to the Air Forces Board for their patience 
during the working out of the initial production grief. 
Such difficulties always arise, even in a program not 
carried forward at the pace of this one. Nevertheless 
under the drive of war it is impossible not to hope 
that each program may be an exception to the rule. 

The principal difficulty was from moisture. The 
terminal board of the transformer proved to be hy¬ 
groscopic, and a new design had to be developed. The 
umbilical switch had inadequate clearance and had 
to be modified. The silver chloride window assembly 
gave trouble, as the presence of grease, dirt, dents, or 



Figure 22. Target at Ocala: at 10,000 ft (left); at 2,000 ft (right). 




DOVE EYE 


65 


irregularities could cause a “built-in signal.’’ Meth¬ 
ods of shop and field inspection had to be developed 
to eliminate defective windows; manufacturing meth¬ 
ods were improved and inspection tightened up. The 
compound used to seal the eye and the window was 
found to be deliquescent. The condition was bad on 
the ground; it was probably worse at an altitude. 

The major source of trouble was associated with 
the excessive humidity in the Orlando area. This is 
spectacularly worse than the atmosphere atTonopah, 
where the preceding tests had been run. It is not 
worse, however, than the test specifications that air¬ 
borne electronic equipments are supposed to meet 
nor, indeed, worse than conditions in the Southwest 
Pacific combat area. Too much emphasis cannot be 
placed on the importance of careful climate testing 
of all components and of the complete assembly be¬ 
fore field trials are attempted. The temptation to 
neglect them is sometimes almost irresistible, but 
succumbing to this temptation can result only in an 
overall loss of time. 

The support of the Air Forces was, however, un¬ 
swerving. In the face of the unsatisfactory tests they 
urged the Division to work out a mass-production 
design and ordered 1,000 units under a transfer of 
funds to NDRC. Accordingly the investigators un¬ 
dertook to select another target which would prove 
satisfactory in Florida summer weather. Further, 
they strongly urged that only production units be 
used in evaluation tests. 

3 6 4 Channel Key Tests 

The target finally selected was Channel Key (Fig¬ 
ure 23), located about one mile northwest of the 
Key West Overseas Highway. While this target was 
relatively weak in comparison with some of the in¬ 
dustrial targets which had now been surveyed (see 
Section 3.8), it was believed that its contrast with 
the surrounding water would make it a satisfactory 
target. Permission for its use was obtained, not with¬ 
out difficulty. 

Only results with the production unit will be re¬ 
ported here. Thirty-one individual releases were 
made. Twenty of these made successful homing 
flights, landing on the target after making observable 
corrections. Of the eleven failures, five were due to 
failure of the bomb shackles to release or to neglect 
on the part of the bombardier to energize the warm¬ 
up circuit to bring the amplifier tubes up to tempera¬ 
ture before release. The difficulty with bomb shackles 


dogged the Division’s entire program and, indeed, is 
said to have been the cause of much bombardment 
inaccuracy in combat. 

One test was of outstanding interest. Eighteen 
Felix bombs were released simultaneously from a 
formation of nine B-17 airplanes flying in loose line 
—abreast in elements of three. Although the average 
spacing of the aircraft in line was supposed to be 400 
ft, the extreme right-wing ship was about 2,500 ft on 
the flank of the target. Figure 24 shows an approxi¬ 
mate map of the trajectories of the eighteen bombs. 



Figure 23. Channel key. 


Fourteen homed well, their impacts forming three 
distinct clusters corresponding to three observed heat 
centers on the target. Figure 25 is a plot showing the 
errors corrected. 

In the closing weeks of the war the Twentieth Air 
Force headquarters placed an order for a squadron 
equipped and manned to take Felix into combat. 
Their assembly was incomplete as the war ended. 

3-7 DOVE EYE 16 

Under urgent representations from the Navy 
Bureau of Ordnance the Division undertook with the 
Polaroid Corporation the development of a quanti¬ 
tative heat-homing scanner. The contractor’s organ- 



66 


FELIX AND DOVE EYE 



Figure 24. Plot of traces of bombs as observed from lead plane. 







DOVE EYE 


67 


ization was eager to develop a complete heat-homing 
missile. The strong belief of the Division, however, 
was that the war was much too far advanced to 
initiate the development of a completely new missile 
system at that time (early 1944). The Navy Bureau 
of Ordnance, however, did sponsor such a develop¬ 
ment with Polaroid Corporation so two projects were 
carried forward cooperatively, the Navy Bureau of 
Ordnance administering work at Polaroid on the mis¬ 
sile structure and its aerodynamic control surfaces 
and servo links, the Division administering the de¬ 
velopment of a quantitative heat-homing scanner 
(Dove Eye). At the close of the war development was 
still incomplete and the project was turned over to 
the Navy Bureau of Ordnance even in advance of 
the preparation of a definitive report. 17 This section, 
therefore, is of necessity general rather than specific 
in treatment. 

From February 1944 on, the Dove Eye was con¬ 
ceived as an infrared detector which would report 
the angular velocity of a line from the bomb to the 
target relative to any stabilized line in space. This 
derivative would be used to effect the minimum aero¬ 
dynamic acceleration necessary to convert a miss into 
a hit. Minimum departure of the missile from its nor¬ 
mal free-fall path would be necessitated, particularly 
if the aerodynamic control system used were con¬ 
tinuous and proportional. 

The first derivative-taking Eye was built and op¬ 
erated during the first three months of the project, 
but because of inherently poor discrimination and 
loss in threshold sensitivity, it was abandoned after 
successful laboratory tests in favor of the present 
lock-on Eye. This “moving-grid” eye obtained the 
derivative by interposing a moving orthogonal grid 
between the target heat flux and the thermistor 
bolometer receptor. Angular velocity of the target 
relative to the Eye showed itself as a frequency shift. 

The final “lock-on” Eye obtained the angular ve¬ 
locity of the line of sight from the bomb to the target 
by continuously pointing along that line. The optical 
scanning system and detector was mounted on the 
rotor housing of a free gyro, with the axis of scan 
collinear with the spin axis of the gyro. Since the 
Eye was made to point at the target by precessing 
the gyroscope, the magnitude of the torque required 
for precession was proportional to the angular veloc¬ 
ity of the line of sight. 

In greater detail, the optical system consisted of a 
2-in. clear aperture spherical mirror with a thermis¬ 
tor bolometer mounted at the focal plane. The mirror 


was displaced horizontally so that its optic axis was 
9.5 mm from, but parallel to, its spin axis (unlike the 
5-degree inclination of the Felix unit). The mirror, 
detector, optical system housing, and initial ampli¬ 
fication stages were mounted on the gyro casing of 
the free gyro, and the mirror was coupled—through 
reduction gearing—to the gyro rotor. The rotary 
motions of both gyro and mirror were coaxial. The 
gyro rotor, gimbal mounted, was precessed in two 
planes by two sets of electric torque motors, one 
motor at each gimbal axis termination. 

During operation the Eye attempted to point at 
the target. If the target was not at the line of sight, 

A 

-N 


DIRECTION 

OF 

FLIGHT 



Figure 25. Errors corrected by Felix in nine-plane 
mission. 


but within the field of view, a signal was initiated, 
its phase in the scan cycle indicating the position of 
the target. The pulses from the detector, after suit¬ 
able amplification and commutation, were used to 
drive the torque motors so as to precess the gyro¬ 
scope and cause the Eye to point at the target. At 
all times the total voltage across the torque motors 
was proportional to the angular velocity of the line 
of sight relative to the fixed line; this voltage, meas¬ 
ured, served also to effect aerodynamic control. 

The heat-sensitive element used in the Dove Eye 
was the thermistor bolometer, initially selected in 
1944 because of its availability, the ease of manufac¬ 
ture of its odd-shaped detectors, its acceptable sen¬ 
sitivity threshold, low microphonics, and low time 
constants. Although suitable evaporated-metal bo¬ 
lometers were produced in connection with the pro¬ 
ject, and successfully used in laboratory lock-on tests 
of production-type Dove Eyes, all the guided mis- 






68 


FELIX AND DOVE EYE 


siles used in homing tests with Dove Eye were pro¬ 
vided with thermistor bolometers. 

The thermistor bolometer used was cruciform, to 
give a nearly circular field of scan. The length of 
each of the four legs of the cross equaled the dis¬ 
placement of the mirror spin axis from its optic axis. 
The ceramic legs of the cross were mounted on a 
quartz backing, and the whole detector hermetically 
sealed in a silver capsule for moisture protection. 
A silver chloride window, suitably filter coated and 
induction welded into the front of the silver capsule, 
admitted desired wavelengths while excluding those 
near and below 1.0 m- 

By June 1946, better than 30 per cent of the Dove 
Eye units dropped from an altitude of 20,000 ft at a 
heated hulk target hit within 50 ft of the target, even 
though the bombsight was set with a fixed error of 
approximately 20 mils. 

38 TARGETS AND TARGET SURVEYS 
3 81 Target Discrimination 

The chief problem in the development of any hom¬ 
ing missile is one of target discrimination. This point 



Figure 26. Industrial target. 


was stressed in Chapter 1 in connection with radar¬ 
homing glide bombs and it is discussed in broad gen¬ 
erality by Dry den in Chapter 12. The problem is par¬ 
ticularly acute in thermal-homing missiles such as 
Felix and Dove. So long as the background surround¬ 
ing the target is uniform, no ambiguity exists. Ther¬ 


mal scanners of any type developed thus far are differ¬ 
ence-measuring devices, but they are sensitive to 
very small differences even though the gross heat 
fluxes compared are large. 

In the Tonopah tests discussed in Section 3.6.2, the 
total thermal flux radiated from the heated target 
was well under 5 per cent of the total flux from the 
field of view. Yet so long as the tests were carried out 
before dawn, when the desert floor was at a uniform 
temperature, troubles with target discrimination 
were minor. After sunrise, however, areas of the ter¬ 
rain warmed at irregular rates, and these irregular¬ 
ities produced signals which masked the desired sig¬ 
nal from the target. 

Marine targets are normally considered to have a 
very uniform background; however, even the ocean 
surface can have thermal gradients of sufficient mag¬ 
nitude to cause difficulties. Military targets on land, 
present a much different aspect. Such targets, for 
example, as that shown in Figure 26, are, in the 
main, industrial and are likely to be surrounded by 
a heavily built-up area. 

The presence of the many discontinuities in the 
thermal array presented by a typical land target has 
the effect of introducing noise into the signal re¬ 
ceived. 

The effect of noise can be minimized, however, by 
integration or smoothing. This corrective measure 
was involved in each of the thermal homing devices 
developed by the Division. In Felix the integration to 
suppress noise was accomplished in the thermal re¬ 
ceptor itself by making it large enough to subtend a 
10-degree field of view. In Dove Eye the same end 
was accomplished by an integrating element in the 
circuitry. The case favoring one of these techniques 
over the other is not clean-cut. Direct thermal inte¬ 
gration as in the Felix manner can be accomplished 
only at the expense of increasing the time constant of 
the bolometer which was already marginally large. 
Circuit integration, the method used in Dove Eye, 
introduces a phase shift in the servo-system loop 
which may be seriously injurious to the accuracy of 
the missile. 

3 8 2 A Target Survey Instrument 18 

In order to determine the suitability of specific 
targets, the Heat Research Laboratory of MIT de¬ 
veloped a quantitative target-evaluation instrument. 
This instrument consisted of a standard Felix scan¬ 
ner mounted in a metal case with a telescope and a 



TARGETS AND TARGET SURVEYS 


69 



GSAP camera. The telescope, the camera, and the 
scanner were all mounted with their axes parallel; 
all had a 10-degree field. The amplifier associated 
with the scanner was in a separate case with the 
necessary batteries. Batteries and amplifier were con¬ 
nected with the surveying head by cables, and the 
surveying head (Figure 27) was provided with a 
gimbal suspension to permit traversing and swinging 
in elevation. 

In addition to photographing the terrain surveyed 
telltale lights connected so as to show the operation 
of Felix rudders and elevators were reflected by par¬ 
tially silvered mirrors onto the film. Further, a 
meter with a logarithmic scale was connected with 
the amplifier so that its scale reading indicated the 
heat flux received in watts per square centimeter; 
this meter was also photographed. The operator ob¬ 
served the meter with one eye while sighting the 


survey head and observing the telltale lights with the 
other eye (see Figure 28). 

This instrument (Figure 29) will evaluate with 
some precision whether any particular target is a 



Figure 28. Typical frame from survey instrument. 


satisfactory objective for an attack with Felix. The 
target-survey report 18 discusses fully the appropriate 
techniques for the use of the instrument. With its 
careful use the successful application of Felix should 
be assured. 






























































































70 


FELIX AND DOVE EYE 


EYE COVER 
CONTROL 


HEAD MASTER 
SWITCH 



MASTER SWITCH 


LIGHT DIMMER 


PILOT LIGHT 


Figure 29A. Surveying head, rear view. 


It is hardly to be expected that a technician with 
an instrument would accompany each combat mis¬ 
sion and advise the commanding officer as to ap¬ 
propriate thermal targets. Rather it is conceived that 
many thermal reconnaissance missions will be flown 
in company with photographic reconnaissance. The 
parallel analysis of these missions, especially if some 
of the photoreconnaissance be in color, will develop 
a target “heat-evaluation lore” so that analysts will 
be able to predict from photoreconnaissance data 


what targets can profitably be attacked with heat¬ 
homing bombs. 

39 FUTURE PROSPECTS 

FOR HEAT-HOMING MISSILES 

The program of the Division in Felix suggests 
further work which may be done in this field. Future 
missiles in the main will probably not be powered by 
gravity; instead, jet propulsion in the transsonic and 







FUTURE PROSPECTS FOR HEAT-HOMING MISSILES 


71 



BOLOMETER 

SPACER 


TRANSFORMER 

UNIT 


Figure 29B. Surveying head, front view. 


supersonic ranges may be employed. Problems of 
higher speed pose new problems in the field of heat 
detection. Scanning speeds must be increased so that 
the distance traversed by the missile during a scan¬ 
ning cycle will remain negligible in comparison with 
the distance remaining so long as additional correc¬ 
tions are likely to be required. This implies the re¬ 
quirement for developing heat receptors of continu¬ 
ously decreasing thermal time constant. The bo¬ 
lometers of the Felix program were small; heat 


detectors of supersonic missiles may have to be 
minuscule. 

Sensitivities, particularly differential sensitivities, 
need to be increased. The relative thermal signal re¬ 
ceived from targets and from a fairly cold back¬ 
ground has been discussed. A better method of 
evaluating the heat from a specific target is needed. 
The comparison of the total heat radiated from an 
area containing the target with a similar area from 
which the target is absent is like measuring the 





72 


FELIX AND DOVE EYE 


weight of a truck driver’s hat by comparing the 
weight of the truck with the driver hatless or covered. 

While the preceding section indicated that promis¬ 
ing heat targets need to be large in area, one pos¬ 
sible exception may exist. In the field of antiaircraft 
guided missiles, particularly in clear weather, the 
contrast between the heat radiated by a jet-propelled 


missile and by the empty sky may well be sufficient 
to produce satisfactory homing control. 

In the entire field, however, no item for further 
study is more important than the exhaustive quanti¬ 
tative survey of all possible targets, including their 
backgrounds, which it may be desirable to attack 
with heat-homing missiles. 



Chapter 4 

ROC AND ITS CONTROLS 


INTRODUCTION 

T he dirigible high-angle bomb in its direct- 
sighted version (Chapter 2) and in its heat-homing 
version (Chapter 3) resulted from a conservative ap¬ 
proach to the guided-missile problem, requiring rela¬ 
tively minor new aerodynamic knowledge or art. The 
glide-bomb program (Chapter 1) was also an exten¬ 
sion of more or less conventional methods to a new 
problem. The Roc program adopted a less inhibited 
view of the problem. It was hoped that thereby a 
broader application of the missile to various tactical 
situations might be attained. 

The original concept was of a radar-controlled 
bomb, the all-weather character of this agency of 
control being most attractive. Later, difficulties in 
radar resolution at the selected approach angles 
forced the decision between revision of the funda¬ 
mental aerodynamic design and the adoption of a 
different control means. In this dilemma the Divi¬ 
sion selected the latter alternative. As the project 
terminated Roc emerged as a television-guided, radio- 
controlled bomb. 

The program was hampered throughout by diffi¬ 
culties in system coordination. The Division’s con¬ 
tractors in the radar and other control fields were 
located on one side of the continent, while the overall 
design of the missile was on the other. This separa¬ 
tion placed a burden on the program which was car¬ 
ried only at the expense of its retardation. Further, 
for the first two and a half years of the program all 
radar development was lodged in a separate division 
of NDRC (see Section 1.7), which was itself heavily 
loaded with programs of high priority so that a 
forthright attack on the radar problems which the 
Roc program faced was impossible. By the time the 
Division had established its own radar group under 
Contract OEMsr-240 (Chapter 1), the promise of suc¬ 
cess with Pelican and with Bat seemed not to warrant 
the diversion of effort of its very limited personnel 
away from the glide-bomb program. 

« GENERAL 1 

The aim of the program was to produce a missile 
whose inherent aerodynamic character provided the 


necessary properties for pilotless control. Only the 
minimum of mechanical devices to correct for aero¬ 
dynamic inadequacy or redundance was permitted. 
Thus the use of an automatic pilot was not only 
avoided but shunned. 

421 Roll Stabilization 

The experience of model-aircraft builders spoke 
strongly for the design of a missile inherently stable 
in roll without assistance of a gyrostabilizer. These 
experimenters, however, used the axis of gravity as 
a fixed reference, and, in general, maneuvers of such 
aircraft were limited to directions not greatly differ¬ 
ing from perpendicularity to gravity. True, loops are 
sometimes executed but at speeds and radii which 
require the net lift of the wing to be positive or dor¬ 
sal. With negative or ventrally directed wing lifts, 
models usually turn over. 

The permissible dynamics of guided-missile flight 
are seriously limited if the control system is con¬ 
strained to use gravity as a reference to prevent roll. 
Guiding in vertical dives then becomes impossible. 
The alternatives are, of course, a gyro system to pre¬ 
serve a fixed frame of reference or else a missile which 
is indifferent to its roll attitude. The latter attack to 
the problem was adopted and led to a design sym¬ 
metrical about any two orthogonal axes normal to 
the axis of flight. Such a structure can develop a lift 
in any transverse direction. From the point of view 
of its ability to maneuver, it is inconsequential what 
roll attitude obtains at any instant. 

4 2 2 Bankless Turn 

The banking of a homing missile in order to correct 
a course error in yaw can create an apparent error in 
pitch. Further, if there is a combined yaw and pitch 
error such that the target appears to be in either of 
the top two quadrants of the field of view (i.e., re¬ 
quiring flattening of the angle of glide) there is a 
positive feedback which can induce sustained oscil¬ 
lations. Such oscillation was not reported by investi¬ 
gators of the Washington Project. This is not a clear 
indication, however, that this effect would be neg¬ 
ligible with another missile. 


73 


74 


ROC AND ITS CONTROLS 


The symmetrical structure indicated as a solution 
for the problem of producing indifference to roll also 
permits turning without banking. Since lift can be 
produced in any direction transverse to the line of 
flight, it follows that moments can be set up and 
sustained about any axis perpendicular to that of 
roll. 

4 2 3 Zero Angle of Attack 

Any aircraft which is to be guided needs a reference 
against which the course may be checked. In ordinary 
airplanes, landmarks and an altimeter serve for con¬ 
tact flying; the compass lubber mark and the horizon 
indicator furnish a first approximation in instrument 
flying. With automatically controlled, pilotless air¬ 
craft, a similar reference is needed—for the control¬ 
ling bombardier if the missile be teledynamically 
guided, for his electromechanical counterpart if the 
missile be fully automatic. 

The usual practice would be to apply an automatic 
pilot to the missile, depending on its gyro mechanism 
to provide a continuing frame of reference. Such prac¬ 
tice violates the basic concept of the Roc program, 
however, which is to make the missile self-sufficient 
without resort to such adjuncts. Self-sufficiency in 
this respect can be achieved by maintaining a refer¬ 
ence axis in the missile continuously tangent to the 
flight path. 

Then, if the control is through a radio link with 
television guiding, the bombardier will be continu¬ 
ously advised, by identifying the center of his re¬ 
ceived picture, of the true direction of flight. The 
picture center does not necessarily indicate the point 
of impact, for in general flight need not be rectilinear. 
An appropriately systematized control needs to be 
worked out. This problem is not easily solved (see 
Section 4.6.2), even with full knowledge of the true 
direction of flight; without such knowledge it is still 
more difficult. 

If automatic homing is applied, the homing system 
can similarly measure the departure of the target 
from the true direction of flight, and an appropriate 
system of control can be designed to apply appro¬ 
priate corrections. Roc was designed to fulfill these 
requirements. In the steady state the requirement 
was well met, the fuselage maintaining tangency to 
the wind stream within 0.5 degree. During transients 
resulting from the application of control there is 
reason to believe that the departure from tangency 
was sometimes significantly greater. In homing 


tests with a photoelectric target seeker (see Section 
4.5.2), the missile appeared to fly with negligible 
angle of attack. In the initial tests with television 
(see Section 4.7.2), there were transient oscillations 
in pitch. The television version (Roc 00-1000) was 
radically different in form from the photoelectrically 
controlled version (Roc-1). In addition, the regime 
of control was changed. Whether the departure of 
the angle of attack from zero during the television 
tests was due to inadvertent loss of aerodynamic 
damping or to feedback from the changed system of 
control remains not fully determined. 

In an automatically controlled missile the dynamic 
or transient conditions are of dominant importance. 
Methods of estimating the transient response in ad¬ 
vance of drop tests are at present incomplete; meth¬ 
ods of measuring it during drops have been incon¬ 
clusive. 

4 2 4 Radar Control 

The original plan for radar control of Roc was to 
have the missile so controlled that it would fly down 
the axis of a microwave beam continuously directed 
from the carrying aircraft toward the target. The 
missile would then ride down the beam. This plan is 
attractive because it places the problem of target 
selection in the hands of a trained operator in the 
airplane instead of requiring the development of an 
elaborate mechanism to make and to maintain that 
selection. It is fundamentally unattractive because 
the sensitivity of control—here defined as the deriva¬ 
tive of field strength with departure from the axis of 
the beam—decreases quadratically as the missile 
falls. Thus the control is least sensitive just before 
impact when it is of crucial importance. 

The objection just cited is outweighed by the ad¬ 
vantages of simplicity in the missile-borne micro- 
wave equipment, which is expendable in quantity. 
There is a further disadvantage, however, which is 
scarcely met by any compensating advantage. Beam- 
borne bombs describe a trajectory exactly similar to 
direct-sighted bombs in the eclipse method of con¬ 
trol if the beam is continuously trained on the target 
(see Chapter 2, Figure 2). Indeed eclipse bombing is 
beam control in the visible electromagnetic spectrum, 
the beam being light originating in the target and 
terminating at the aircraft. The same arguments 
which spoke against its use for the dirigible high- 
angle bomb speak with nearly equal force here. The 
only mitigating feature in this instance is the pres- 



GENERAL 


75 


ence of reasonably large supporting surfaces, with 
corresponding reduction in the minimum attainable 
radius of trajectory curvature. However, the ability 
of a radar operator to keep a guiding beam continu¬ 
ously on the target is doubtful. Wavering of the beam 
when the missile is near the end of its flight would 
give rise to intolerably large requirements of trans¬ 
verse lift. The decision was made, therefore, to design 
Roc as a radar-homing missile. The radar was to be 
of substantially the same type as for RHB-Pelican 
with illumination provided either from the parent 
plane or from another in the squadron. 

To make the corrections to the flight path non- 
oscillatory, the rate of change of transverse lift was 
made proportional to error heading and to its first 
time derivative. 



Figure 1. Roc-1. 


The initial experiments to prove the homing char¬ 
acteristics of the system were made against a strong- 
light source with an improvised photoelectric target 
seeker. The tests were supposed to be with a full- 
sized (although not necessarily a fully loaded) mis¬ 
sile, which could be the prototype of a combat 
weapon (Figure 1). However, by the time the tests 
had been completed, the earlier requirements of the 
Services for a design utilizing a 14-in. diameter bomb 
had changed, and the new requirement was for one 
of 18%-in. diameter. 

In addition, while the experimental program with 
photoelectric target seekers was proceeding, crucial 
experiments with radar had shown that reliable hom¬ 
ing signals could not be obtained with angles of ap¬ 
proach steeper than 50 degrees if the source of radia¬ 


tion illuminated the target at an angle steeper than 
11 degrees, both angles being referred to the hori¬ 
zontal. 2 Roc had been designed to fly within a vertical 
cone having an apex angle of 90 degrees located on 
the target. The margin between the flattest glide 
path for Roc (approximately 45 degrees) and the 
steepest angle for radar homing (50 degrees) left 
little margin for flexibility, either in design or in use. 
The requirement of illumination at a flat angle (11 
degrees) implied a two-airplane maneuver for attack. 
The Services found this limitation intolerable. 

The requirement of the new payload—the 1,000-lb 
GP bomb instead of the 14-in. SAP bomb—meant a 
new design with considerable increase in wing span. 
Problems of carriage were considerably increased, 
especially as the Navy insisted that the missile be 
capable of use with carrier-based aircraft. 

4 2 5 Television and Radio Control 

Faced with a new design on account of problems 
of stowage with the 1,000-lb GP payload, the re¬ 
quirement of use with small aircraft, and the ap¬ 
proach-angle restriction previously unknown in the 
use of radar for homing, the Division had three 
choices: (1) in revising the design to accommodate 
the 1,000-lb GP bomb in reduced space; aerodynamic 
changes could be made which would so flatten the 
trajectory as to make radar homing practicable with 



Figure 2. Roc 00-1000. 

single-airplane tactics; (2) the aerodynamic perform¬ 
ance could be left unaltered and a new means of con¬ 
trol sought; or (3) the project could be terminated. 

While calculations showed that minor revisions of 
the design would produce an angle of approach flat 
enough to accommodate radar, this change involved 
removal of the dive brakes, and it was not known by 
what extent this would influence the missile in its 
property of flying continuously tangent to the flight 






76 


ROC AND ITS CONTROLS 


path. The image orthicon had been developed (see 
Chapter 5), and it appeared likely that a miniature 
version of small enough compass to fit within Roc 
could be developed. Accordingly, the contractor ac¬ 
cepted the Division’s request to continue the project 
toward television control. 

Furthermore, the increased maneuverability of 
Roc (Figure 2) as compared with Razon—8,000-ft 
minimum radius of curvature compared with some 



Figure 3. Wing assembly for Roc-1 looking radially 
outwards. 

20,000-ft—gave promise of successful application of 
Roc to the eclipse technique of direct-sight control. 
This was strongly urged as an interim measure pend¬ 
ing the development of Mimo (miniature image or¬ 
thicon camera and transmitter). Such a program was 
particularly indicated in view of the fact that while 
the development of the Crab attachment (see Section 
2.6) had been started, its accuracy had not been es¬ 
tablished. An eclipse technique gives exact accuracy, 
but it is most difficult to achieve. 

As the war closed, tests with television were under 
way. The direct-sighted version had been abandoned 


in view of the marked success with Razon when used 
with Crab and especially with the addition of Jag. 
Further investigations with Roc and Mimo are to 
continue under the supervision of the Air Technical 
Service Command. 

43 ROC SYSTEM DYNAMICS 3 

The requirement that Roc should be capable of 
generating lift in any direction and of flying with 
zero angle of attack led to the embodiment of Figure 
1. Cast magnesium wings were constructed with a 
quadrant yoke at the root such that, when they were 
bolted together, they formed the central portion of 
the fuselage. The wing was hollow, providing room 



Figure 4. Two-dimensional projection of trajectory. 


(Figure 3) for a motor and gear box to drive the lift¬ 
generating flaps, which extended the full semispan 
of the wing. 

The establishment of a law to govern the control- 
flap position is the outcome of the solution of the 
problem of system dynamics. Six degrees of freedom 
are involved, and the relationships are generally non¬ 
linear. A simplified analysis is, however, rewarding. 
In his analysis of longitudinal stability of glide 
bombs (see Section 1.4.1), Skramstad suggested a 
control regime defined by 

5 = kid + k 2 d (1) 

This contractor used a method which produced flap 
operation according to: 

5 = faO + M (2) 

There are many control methods possible which will 
yield the desired result, namely approach toward the 















ROC SYSTEM DYNAMICS 


77 


target along a pursuit curve without superposed os¬ 
cillations. 

Consider now the two-dimensional simplification 
of Figure 4. The missile is at A with a velocity V and 
a course error 6 which would result in an expected 
miss h. A lift L must be generated to reduce h to zero 
and in a nonoscillatory manner. (Of course it may be 
impossible to reduce h to zero, a finite lower limit 
existing for the radius of trajectory curvature R'. 
That phase of the problem does not concern us here.) 

The lift L is balanced by dynamic reactions due to 
imparting a centripetal acceleration to the missile 
and an angular acceleration about an axis perpen¬ 
dicular to the flight-path tangent and to the lift L. 

L = amd + bid (3) 


where m is the mass of the missile; 

I is the moment of inertia of the missile; 
a is a constant of proportionality; 
b is an assumed constant relating the restor¬ 
ing moment to the angle of attack. 


If 


amd > bid 


(4) 


then the missile can be deemed to behave like a mas¬ 
sive particle. Further, if linearity between 8 and L 
is assumed, we can write from equation (1): 


L = kid + k 2 d 

dO 

amd = kid + k 2 d 

(5) 

d(k 2 — am) + kid = 0 

(6) 

kit/{k 2 - am) 

U — Uq6 

(7) 


which is decadent if k 2 < am. It satisfies our require¬ 
ment if d is small so that 


Similarly with equation (2): 

L = kzd + h A d (2') 

amd — kzd + k\d 



am am 


whence 

0 o = 006 k “ lam sin ( |/— - 44) t ( 8 ) 

\\ am a 2 m 2 / 

Thus, in the simplified version at least, either flap- 
control law will yield stable approach to the target 
with judicious selection of the constants. 

In the more complete analysis where angular ac¬ 
celeration in pitch or yaw is not considered negligible, 
control of the lift by equation (!') can lead to damped 


oscillations similar to equation (8). Control of the 
rate of change of lift in accordance with equation 
(2') results in a cubic equation for the roots. The real 
root can be made degenerative; the other roots are, 
by proper choice of constants, conjugate and repre¬ 
sent a damped oscillation. 

The control law of equation (1) thus seems to lead 
more directly to a stable trajectory than does that of 
equation (2). However, it also leads to greater driving 
power for the control flaps. The abrupt appearance 
of a course error calls for the immediate appearance 
of lift, with a correspondingly large expenditure of 
power to overcome the hinge moments at a high rate. 
Further, while it is impossible to attain instantane¬ 
ously the flap velocity 8 of equation (2) required by 
the sudden appearance of an error in heading or rate 
of departure, it is one derivative easier than the in¬ 
stantaneous attainment of a flap angle 8. In sum¬ 
mary, it would appear that one system of control 
leads more readily to the desired flight path but is 
much more difficult to mechanize. The implication is 
that time lags will force a greater departure from the 
idealized program in the case of equation (1) than 
in the case of equation (2). 

One feels that a full exploration should be made of 
the control possibilities implicit in 

n = °o m = » 

L k n p n (8) = £ k m p m (d) (9) 

n= — * m= — oo 

In equation (9) p is the time derivative operator de¬ 
fined by: 



Negative values of n and m indicate integrals. 4 The 
generalized exploration of equation (9) with the quan¬ 
titative rejection of such terms as are redundant or 
insignificant might well teach where the best engin¬ 
eering compromise in design of a control system lies. 
Its generality is extremely suggestive of power in the 
approach to other missile problems. 

It is significant in this connection that in develop¬ 
ing controls for the V-2, the Germans 5 used a control 
regime defined by 

coo L -f- c c\L -f- c o 2 L = a' Edt 

-f- coo'tfo Oil E + c 0 2 E (11) 

where E is a measurable error function. Moreover, 
the parameters c o n and 00 / were not necessarily con¬ 
stant but were permitted dependence upon error or 
its correction. 




78 


ROC AND ITS CONTROLS 




Figure 5. Exploded cutaway and views of Roc-1. 


















MISSILE DESIGN 


79 


4 * MISSILE DESIGN 

441 Cross-Wing Roc Structure 6 

The final design of the cross-winged version is 
shown in Figure 5. All fixed members are castings. 
The wings and central fuselage section are of mag¬ 
nesium; the forebody, afterbody, and empennage are 
of aluminum or of zinc alloy, depending on the weight 
of missile desired. In the combat version it was 
planned that all castings would be magnesium. 

The space forward of and surrounding the nose of 
the payload was planned to house the radar-homing 
equipment; space abaft the payload was provided for 
the power supply. In the test birds no payload was 
used, so that extreme effort to compress the control 
into minimum compass was not exerted. 

The principal dimensions are given in the following 
tabulation. 


Length overall (in.) 

Wing span (in.) 

Effective area of 2 wings (sq ft) 

Weight with payload (lb) 

Wing loading (psf) 

Test missiles 

Weight without payload (lb) 

Wing loading (psf) 

Flap hinge line (per cent of wing chord) 
Tab hinge line (per cent of wing chord) 
Wing system section 
Root 
Tip 

Tail-fin span (in.) 

Tail-fin area of 2 fins (sq ft) 

Tail-fin section 
Foot 
Tip 


95.98 

77.0 

10.95 

1,300 

119 


500-600 

45-55 

55 

9 

NACA #0011 
NACA #0009 
40 
1.8 

NACA #0009 
NACA #007 


The flap is partially balanced, its rounded nose 
fairing into a recess in the fixed wing placing the 
hinge line at 55 per cent of the section. The tab at 
the trailing edge of the flap gives approximately full 
balance. With the flaps in the neutral position the 
wing is without camber, and its chord is parallel to 
the fuselage axis. The tail-fin chord is continuously 
parallel to the fuselage axis. The empennage is fixed 
but it is so constructed that the planes of the fins can 
lie either in or midway between the planes of the 
wings. The latter configuration gives somewhat bet¬ 
ter trim stability. 

The full-sized missile was tested in the wind tunnel 
at the Ames Aeronautical Laboratory, Moffett Field, 
California. 7 This test showed that the missile would 
fly with substantially zero angle of attack, that a 
maximum transverse lift of 2,000 lb could be gen¬ 


erated in the plane of two wings at 400 mph or of 
2,800 lb in a plane midway between two wings. The 
corresponding minimum radii of curvature with the 
fully loaded missile are 6,940 ft and 4,950 ft respec¬ 
tively. The brakes hold the terminal velocity to 
about 500 mph. Rolling torques due to casual screw 
asymmetries are negligible! and are readily killed by 
differential operation of one pair of flaps. 

The flaps are made of molded resin-impregnated 
Fiberglas in stressed-skin hollow construction. They 
are hinged on bushed pins at the root and tip. A 
pin at midspan is supported on open slots to prevent 
binding due to load deformation of the wing and flap. 

The balancing tabs are of molded resin-impreg¬ 
nated veneer. They are supported on knife edges rid¬ 
ing in molded saddles, being held in place by the 
cables that control their angular position with respect 
to the fixed position of the wing and to the flaps. 

The drive for the flap (Figure 6) is an electric mo¬ 
tor and gear box, one assembly for each wing. The 
overall gear and belt reduction, 7,582 to 1, yields a 
flap speed of 6 degrees per second with a motor speed 
of 7,582 rpm. The input to the motor is approximately 
8 watts, and the overall efficiency of the system is 
approximately 25 per cent. 

The tab is driven through a cable belt. The belt 
is fastened to a pulley fixed to the wing and con¬ 
centric with the hinge line of the flap. It is also fast¬ 
ened to a pulley, which is fixed to the tab concentric 
with its hinge line. Pulley ratios are such that the 
angular motion of the tab with respect to the wing 
is 20 per cent of the angular motion of the flap. The 
angle between the flap and the tab is thus 80 per 
cent of the wing-flap angle. A slight loss in lift due 
to the tab is the price for substantial elimination of 
the hinge moment. 

4 4 2 Control System 

The basic purpose of the program was to develop 
a radar-homing missile. In the absence of expendable 
radar equipment for tests, a photoelectric target 
seeker was improvised. This target seeker, which is 
described more fully in Chapter 9, was designed to 
give an output closely similar to that expected from 
the radar equipment. For small errors in course, the 
output of the target seeker is a voltage nearly pro¬ 
portional to error in heading (Figure 7); thereafter 
it holds nearly constant until the limit of the field of 
view is approached. For small errors, then, the volt¬ 
age is proportional to error heading. In order to 



80 


ROC AND ITS CONTROLS 



Figure 6. Schematic diagram of flap and tab. 



















































MISSILE DESIGN 


81 


apply the control regime selected as a result of the 
arguments in Section 4.3, this voltage must be made 
to drive the control flaps at a speed proportional to 
it and to its first time derivative. Then 


since 


5 = k\e + k 2 e 


e = c6 

for small values of 0. 

Pacific Division of Bendix Aviation Corporation 
under Contract OEMsr-1002, 8 developed an amplifier 
to perform these operations. A preamplifier driven by 
the photoelectric scanner produces a voltage in a two- 



40 30 20 10 0 10 20 30 40 


TARGET LEFT IN DEGREES TARGET RIGHT IN DEGREES 


Figure 7. Expected output of radar equipments; 
actual output of Douglas PE ta'rget seeker. 


channel output such that the voltage level in either 
channel is zero or 


0 < ci = cd 

This voltage is applied to the network of Figure 8. 
Now if the attenuation of the network is made such 
that the output e 2 is small compared to the input e h 
then the output of either channel is either zero or 

0 < e 2 = CiCi -f c 2 e i 

It is not necessary to establish any specific values of 
Ci and c 2 since the level of the voltage e 2 can be ad¬ 
justed in succeeding stages. Their ratio must, how¬ 
ever, be established here since it is this ratio that 
determines the nonoscillating quality desired for 
Roc’s flight. This ratio is determined by R according 
to the relation 


8fci /h ~ 5.5 v / 6-25(A:i//b 2 ) 2 - 33/ci /k 2 + 30.25 
H ~ 2 - 2.lki/k 2 

In this evaluation ki and k 2 are the constants of pro¬ 
portionality in the fundamental Roc control equa¬ 
tion. For nonoscillatory approach to the target, the 
ratio k\/k 2 should be not greater than 0.2. This re¬ 
quirement results in a value of 0.625 megohm for R 
in the circuit of Figure 8. 

A voltage satisfactorily proportional to error in 
heading and to its first derivative having been de¬ 
veloped, it is necessary to make the flap motors fol¬ 
low this voltage in speed, irrespective of load torque. 
The output of the differentiator-mixer (approximate¬ 
ly 0.3 v maximum) is amplified to about 150 v maxi¬ 
mum through two stages of push-pull 6SL7-GT in a 
Class A circuit. This voltage triggers thyratrons to 
control the flap motors (Figure 9). 

The flap motors are separately excited at 12 v. 
The armature is especially wound for 100-v d-c aver¬ 
age when supplied from a half-wave rectified source. 
The power supply is a 100-c vibrator and trans- 



Figure 8. Differentiating-mixer network. 


former. During the conducting half-cycle, the thyra- 
tron will fire if the grid is at a sufficient potential to 
trigger it. The bias at the instant of firing is deter¬ 
mined by the grid voltage supplied by the d-c am¬ 
plifier and the cathode voltage. Since at this instant 
no current is flowing, the cathode voltage is equal to 
the voltage generated by the motor armature and 
(the motor being separately excited) exactly propor¬ 
tional to its speed. Thus for any grid voltage there is 
a motor speed which will just prevent firing of the 
thyratron. At any load torque a sufficient percentage 
of the half-cycles will pass current to develop the 
























































82 


ROC AND ITS CONTROLS 


torque at the speed demanded by the grid voltage 
(Figure 10). Thus the extremely flat speed-load char¬ 
acteristic is achieved. It should be noted that this 
application imposes a severe duty on the motor. The 
entire armature circuit floats at an average voltage 
determined by leakage resistances; it is subject to 
violent surges when the tubes fire or cut off. It was 
necessary to insert a 0.001-juf capacitor between the 
thyratron grid and cathode to prevent surges from 
becoming regenerative. 



Figure 9. Thyratron motor-control circuit developed 
by Radiation Laboratory, MIT. 


One pair of wings is electrically locked in step. This 
is accomplished by feeding into the d-c amplifier 
controlling each flap motor a voltage proportional 
to the difference between its travel and that of its 
mate. Thus, if one flap gets ahead of the other, its 
speed is immediately reduced while that of its partner 
is increased. This circuit maintains identity of the 
flap displacement within 1 degree—approximately 
4.3 per cent of full travel. 

The other pair of flaps are purposely allowed to 
operate separately. A gyro sensitive to rate of roll 
biases their speed of operation differentially, intro¬ 
ducing a differential flap speed of 0.4 degree per sec¬ 
ond if the rate of roll of the missile is between 0.833 
rpm and 2.5 rpm, and a differential speed of 1.6 de¬ 
gree per second if the rate is in excess of 2.5 rpm. 
Figure 11 shows the electronic control as used for the 
photoelectric target seeker and as planned for use 
with radar. 


4 5 TESTS WITH PHOTOELECTRIC 
HOMING 

Preliminary Tests 

As a proof of the nonoscillatory character of the con¬ 
trol regime, a scanning head and the associated elec¬ 
tronic apparatus were mounted on a truck and made 
to home on a strong light.The driving was done on a 
dry lake, and the driver turned the wheel in such a 
manner as to match the motions of a flap position 
indicator controlled by the scanner and amplifier. A 
control factor for the truck can be defined as the 
product of the wheel displacement and the corre¬ 
sponding turning radius. For a given air density a 


11,000 

10 non 

\ 

.r 

Ml" 

THYRATRONS FIRE 

1 1 1 1 

EVERY HALF-CYCLE 


9000 

*\ 

< 












>5V- T 

HYRAT 

RON B 

IAS 




7000 

6000 




\ 










\ 

N 4! 

jV-TFn 

r'RATRC 

)N BIA! 

S 


5000 

4000 






X 

s. 










V 

V 



3000 

POOO 



22 ' 

/- TH1 

'RATRO 

N BIAS 
















1000 

0 






















0 20 40 60 80 100 120 140 160 180 200 

PER CENT OF FULL LOAD TORQUE 


Figure 10. Motor speed versus load torques and bias. 

similar product of flap angle and turning radius is 
descriptive of Roc. The ratio of these products is the 
scale factor by which the driver should relate steer¬ 
ing-wheel motion to flap-indicator motion. 

This relationship was approximately adhered to. 
In eight tests the truck homed on the light along a 
nonoscillatory path. In other tests the differentiator 
was cut out so that the steering-wheel speed was 
made proportional to error-heading only. In these 
tests the course was oscillatory and soon got out of 
control. 

Prior to making drop tests, all apparatus was sub¬ 
jected to altitude and temperature tests. In addition, 
a test was devised to give a quantitative evaluation 
of the control system. A small light was mounted on 






























































TESTS WITH PHOTOELECTRIC HOMING 


83 


a wheel located in front of the missile and rotating 
about an extension of its roll axis. Thus, with con¬ 
stant wheel speed the motion in the yaw and pitch 
senses was simple harmonic, with all its derivatives 
well known. With small angles of parallax subtended 
by the wheel diameter, the response of the flaps could 
be quantitatively checked; for larger excursions of 
the light, a more involved analysis was required. 



Figure II. Bendix motor-control amplifier. 

4 5 2 Drop Tests 9 » 10 

In spite of the careful preliminary testing, five of 
the nine missiles dropped failed for various reasons. 
Drops No. 2, 3, 4, and 5 were successful, the first 
having failed through malfunction of the arming 
switch. This uniformity gave rise to an optimism 
which was perhaps unfounded. The program was 
then plagued by a variety of minor but frustrating 
difficulties, and the next four drops failed. This re¬ 
versal gave rise to a pessimism which was equally 
false. 

The tests were carried out at the rocket range of 
California Institute of Technology on Goldstone 
Lake near Barstow, California. The target consisted 
of an array of 100 battery-powered photoflood lamps 


supplemented by pyrotechnic flares. The light inten¬ 
sity at the release point, 10,000 ft above the target 
and average 25-degree lead angle, was approximately 
10~ 2 footcandle. The flux entering the scanning eye 
as the target entered the field of view was estimated 
at 300 microlumens, approximately 10 times the 
threshold for operation of the target seeker. 

The drops were fully instrumented (see Chapter 
8). Misses of 260, 96, 68, and 30 ft were scored. Fig¬ 
ures 12 and 13 present the frontal and profile traces 
of the trajectory of a typical drop as recorded by 
cameras with open shutters set up approximately 



Figure 12. Frontal trace of PE homing drop No. 4. 


4,000 ft from the target in the plane of approach and 
2,000 ft on the flank. A more precise determination 
of the trajectories was made from the analysis of two 
phototheodolite records taken from the ends of a 
carefully surveyed base line. Figure 14 is a recon¬ 
struction of the plan and side elevation of the trajec¬ 
tory shown in the preceding photographs. 

The trajectory shown in the photographs of Fig¬ 
ures 12 and 13 and in the plots of Figure 14 shed in¬ 
teresting light on the dynamics of Roc control. For 
small errors of heading, the target seeker gives a pro¬ 
portional output. In the azimuthal sense the error in 
heading was small, and the control response was 






84 


ROC AND ITS CONTROLS 


quantitatively sensitive to it and to its first time de¬ 
rivative. The frontal trace (see Figure 12) of the 
trajectory reflects this condition, which resulted in 
non-oscillatory homing. 

In the range sense, however, the error is not always 
small. It starts, of course, at the complement of the 
lead angle to the target, corrected for the attitude of 
the airplane at the instant of release. For this drop 
the lead angle was 19 degrees; the correction for atti¬ 
tude of the airplane was negligible. The initial range 
error in heading was thus 71 degrees, which lies be¬ 
yond the field of view of the target seeker. As the 



Figure 13. Profile trace of PE homing drop No. 4. 


target enters the field of view at a range error in 
heading of about 33 degrees (see Figure 6), the out¬ 
put is in the correct sense but its slope is wrong, 
producing an increasing signal for a decreasing angle. 
This results in negative damping until constant out¬ 
put is reached at approximately 21 degrees error. The 
damping disappears with constant output and only 
appears in the correct sense after the error has been 
reduced to 10 degrees. The profile trace (Figure 13) 
and the side elevation plot of the trajectory (Figure 
14) show the resulting overshoot in the initial portion 
of the flight caused by negative damping during the 
early phases of the oscillation. The remaining portion 


of the trajectory after the first swing was well 
damped. 

The experiments were further instrumented by 
motion-picture cameras mounted on the empennage 
and pointed forward to record the view “seen” by 
the scanner. Tangents drawn to the trajectory pro- 



5000 4000 3000 2000 1000 0 

RANGE FROM TARGET IN FEET 


Figure 14. Plots of PE homing drop No. 4 from 
phototheodolite records. 


jections derived from the phototheodolite records 
disclose the instantaneous direction of the missile’s 
flight. The center of the corresponding motion-picture 
frame shows the spot on the terrain toward which the 
missile was pointing. A comparison of these data 
showed that Roc-1 flies with small angle of attack. 



























DESIGN OF ROC 00-1000 


85 


This conclusion can hardly be said to be more than 
qualitatively established. Some of the motion-picture 
film was lost at impact. In some of the drops one or 
both of the phototheodolite operators were unable to 
track the missile. Finally the graphical differentiation 
of an empirical curve to establish the direction of its 
tangent is inherently not very accurate. 

In addition to the series of nine tests just described, 
three drops were made of Roc missiles equipped with 
a photoelectric target seeker developed under Con¬ 
tract OEMsr-1182 by Fairchild Camera and Instru¬ 
ment Company. 11 These drops, marred by flare and 
gyroscope failures, were inconclusive. The target 
seeker is described in Chapter 9. 

The problem of coordinating the activities of an 
aircraft crew, two phototheodolite stations, the tar¬ 
get operation, several camera stations, and the usual 
observers from the Division and from the military 
is not to be underestimated. This experience of the 
Division’s contractor reinforces the conclusion drawn 
from each of the other programs. An adequate ex¬ 
perimental range equipped with good communica¬ 
tion, adequate transport, and facilities for the routine 
collection of formal quantitative records is essential 
if success in this field is to be quickly attained. 

4 6 DESIGN OF ROC 00-1000 1213 

461 Structural Design 

The tests described in the preceding section dem¬ 
onstrated with some authority that the Roc concep¬ 
tion resulted in a system of missile and control which 
would produce satisfactory homing flight. However, 
the wing span (77 in.) precludes carriage within the 
bomb bay; with external carriage there is inadequate 
clearance on carrier-based aircraft; and finally, the 
original payload contemplated (the 14-in., 1,100-lb 
SAP bomb) was considered an undesirable weapon 
for the bulk of the remaining targets to be attacked. 
Accordingly, the Division requested the contractor 
to undertake the development of a missile for carrier- 
based aircraft. The payload was to be the 1,000-lb 
GP bomb, and the method of carriage was to permit 
landing with the missile unexpended on the deck of 
an aircraft carrier in the event of an abortive mission. 

The resulting design eliminated the cruciform 
wings with adjustable camber and the cruciform em¬ 
pennage. In the new design a cylindrical shroud re¬ 
places each (Figure 15). The wing shroud is Cardan- 
mounted so that it can be rocked to produce lift in 


any direction normal to the roll axis. The stabilizer 
cylinder is fixed. Within the stabilizer ring two aile¬ 
rons were planned to limit the rate of roll; in the final 
version, however, four ailerons were provided, so 
controlled as to eliminate roll. 

The principal dimensions and the performance 
constants as determined from wind-tunnel tests 14 are 
compared with those of the cruciform Roc-1 in the 
following table. 



Roc-1 

Roc 


Combat Version 

00-1000 

Weight complete (lb) 

1,300 

1,662 

Payload (lb) 

1,100 

1,000 

Turning radius (ft) 

6,950 

7,500 

Wing span (in.) 

77 

48 

Wing area (sq ft) 

10.95 

8.86 

Wing loading (psf) 

119 

188 

Length overall (in.) 

96 

148 

Terminal velocity (mph) 

450 

400 

Mean wing chord (in.) 

20 

30 

Tail span (in.) 

40 

30.6 

Mean empennage chord 
(in.) 

8.3 

5.3 

Ailerons 

None 

4 

Time to apply full control 
from neutral (seconds) 

2.5 

1.6 

Max. hinge moment main 
control surfaces (lb-in.) 

* 

2,170 

Wing airfoil section 

NACA #0011-Root 

Douglas N 8 

Stabilizer airfoil section 

NACA #0009-Tip 
NACA #0009-Root 

Douglas 


NACA #0007-Tip 

81080-18 

Aileron airfoil section 

None 

Douglas 


81080-18 


*Balanced by tab. 

With the exception of a false nose, the entire con¬ 
struction is of aluminum-alloy sheet, with the bomb 
itself forming the central portion of the fuselage. A 
plastic nose section, fastened to the bomb through 
the nose-fuze threads, houses the television camera. 
Abaft the bomb a fuselage structure carries the re¬ 
maining electronic equipment, the power supply, one 
wing actuator, and the gyro-aileron system. 

The wing is of monospar construction with the 
major portion of the training section permanently 
attached to the spar. Removal of the nose section 
and a generous panel of the trailing section yields 
access to wiring, strut bearing, and the second wing- 
actuators. 

The wing is carried by a strut (Figure 16) which is 
hinged in the fuselage to rock the wing in yaw. A 
bearing support between the wing and the strut pro¬ 
vides freedom for rotation of the wing in pitch. Thus 
lift and drag are carried by the strut in bending; 
hinge moments in pitch are carried axially and hinge 
moments in yaw are carried in bending. The strut is 




86 


ROC AND ITS CONTROLS 


STABILIZER 


DIVE BRAKE 
ANO ANTENNA 


WING 
ACTUATOR! 


NOSE 



AILERONS 


AILERON - ORIVE 
SOLENOIDS 


.TELEVISION 

CAMERA 


STRUT 


AFTER FUSELAGE 
HOUSES POWER 
SUPPLY, RADIO 
RECEIVER, AND 
CONTROL AMPLIFIER 


Figure 15. Cutaway view of Roc 00-1000. 


fabricated of X4130 steel, heat-treated to 125,000- 
145,000 psi to withstand the foregoing design loads, 
calculated at 91,200 psi at the critical section. 

In the cruciform version (Roc-1) hinge moments 
of the control flaps are nearly eliminated by balanc¬ 
ing tabs. With a movable cylindrical wing this tech¬ 
nique is not applicable. 



Figure 16. Wing strut, Roc 00-1000. 


The study to determine the basis of the control 
system was not complete as soon as the missile de¬ 
sign was ready for its findings. There is some argu¬ 
ment, in the case of television, for using a quasi¬ 
homing control in which the controlling bombardier 
tracks the received image of the target and in so 
doing automatically transmits a signal to the missile 
proportional to the error in heading and to its first 
time derivative. The possibility that a control system 
based on the angular velocity of wing incidence, 8 , 
might be required, rather than a system based on 8 , 
spoke strongly for using the same electronic com¬ 
puter developed for Roc-1. 

The large hinge moments and the use of a single 
wing, however, imply a much heavier load on the con¬ 
trol motor than is the case in Roc-1, the entire power 
supply for which is passed through a single thyratron 
rectifier. It became necessary, therefore, to exert ex¬ 
traordinary effort toward high efficiency in the wing 
actuators. A ball-bearing, thread-nut drive was de¬ 
veloped by Western Gear Works having a mechanical 
efficiency of approximately 80 per cent and an overall 
efficiency including the motor of about 40 per cent. 

The decision of the Division to attempt direct-sight 
control as an interim means of control, pending the 
development of a suitable television or automatic 









DESIGN OF ROC 00-1000 


87 


homing equipment, made it necessary to develop ab¬ 
solute roll stabilization. This the contractor agreed 
to undertake although the extrapolation of the ex¬ 
perience with Azon to Roc seemed hazardous in view 
of the greatly increased maneuvers to be required. 
The same gyroscope assembly developed for Azon 
and Razon was applied to the aileron control. Air¬ 
craft practice suggested a modulating control for the 
ailerons such that small excursions from a mean posi¬ 
tion would provide transient roll stability. With the 
on-off contact arrangement provided from the free 
gyro, such a control with motor-driven actuators 
proved wholly successful. A solenoid drive capable of 
following gyro oscillations up to 5 c was used. 

4 6 2 Control System 

Control with direct sight or television implies a 
radio-control link. At the outset of this phase of the 
program it was not clear what mode of control would 
be most appropriate. For television, as just explained, 
the contractor favored the same regime used for au¬ 
tomatic homings, 

8 = kid “b k 2 d 

the bombardier keeping a reticle superposed on the 
televised image of the target. The Division felt that 
arguments such as temporary blocking of the radio 
signal spoke strongly in favor of a regime expressed 
by: 

<$ = kid + k 2 d 

Moreover, in view of the success with Azon, which 
used on-off control, there was strong compulsion to 
experiment with that system for Roc. 

The Division worked closely with Section 7.2 (Air¬ 
borne Fire Control), which recommended a computer 
which would steer the missile along a nearly straight 
interception course rather than along a pursuit curve. 
Simulative studies 15 with Division 7’s contractor at 
Columbia University disclosed that for the direct- 
sighted version a link which provided wing deflection 
proportional to control-stick motion is adequate. For 
the television version, Douglas constructed a model 
range with a test cart carrying television and a radio 
receiver. This model reproduced the missile flight in 
two coordinates and time. Many modes of control 
were explored. 10 With the aid of this range and the co¬ 
operation of Section 7.2, a computer was developed 
which resulted in an approximation of a collision 
course. 


The computer projects on the television screen a 
reticle which moves across the screen at a speed pro¬ 
portional to the wing-incidence angle 8 and to the 
actual reticle position. The actual equation of motion 
of the reticle is: 

^=f^ + f(x»-x+|°) 

Where X = reticle position in degrees with respect 
to the center of the television field of 
view, 

X 0 = target and reticle position at the time 
computer is started, 

y 0 = angle in degrees between gravity vector 
and trajectory of missile at the time the 
computer is started, 

K = (S/2m)(dC l/(18), the aerodynamic re¬ 
sponse coefficient of the Roc missile, 
p = the average air density during the ef¬ 
fective phase of guided drop, 
v = average value of the missile velocity, 
g — acceleration of gravity. 

In a normal television drop the target appears at 
the bottom of the screen and moves upwards across 
it as the missile noses over. The reticle is located near 
the bottom of the screen. As the target image crosses 
the reticle, the computer is turned on, which starts 
the reticle moving up screen in accordance with the 
above equation. The operator then so manipulates 
the control stick as to keep the target on the reticle. 
When the appropriate lead angle is established, the 
motion of both target and reticle stops. 

Various control regimes will satisfy the computer 
equation. It is necessary only that the control stick 
feed into the computer a value proportional to the 
wing incidence attained. With the aid of this com¬ 
puter, runs on the model range showed a tenfold im¬ 
provement in accuracy (12 ft average miss as com¬ 
pared with 120 ft) over unaided guiding with tele¬ 
vision alone. 

As the project closed, it appeared that the com¬ 
puter renders equally accurate an on-off system of 
control in which the wing-incidence angle is propor¬ 
tional to the net total time of control, right-and-left 
control time being taken in opposite sense. This sys¬ 
tem is much the simplest of those that have any hope 
of success. It was not tried in drops made under Divi¬ 
sion sponsorship, although five missiles have been 
built for this control regime. 

It is obvious from the foregoing that it was not 
until the project terminated that the optimum mode 



88 


ROC AND ITS CONTROLS 


of control which determined the control-link design 
was established. It appeared obvious that a quantita¬ 
tive radio link would be required, delivering a signal 
proportional to control-stick deflection. The decision 
as to whether the wing-incidence angle should be 
made proportional to the received signal, to the con¬ 
trol signal, and to its time derivative, or whether the 
angular velocity of wing incidence should be con¬ 
trolled, had to be held in abeyance. 

The decision having been made to use the same 
standard transmitter that controlled Azon and Ra- 
zon, Bendix developed a proportional control link 
based on it. Two versions appeared. In the first, the 
carrier which operated in the 100-mc band was conti¬ 
nuously modulated by four audio tones. The differ¬ 
ence between the degree of modulation of tones 1 and 
2 gave the amount of “right” control—"left” if tone 
2 was greater than tone 1; similarly the difference be¬ 
tween tones 3 and 4 gave quantitative control in 
range. The signals were received, demodulated, recti- 



Figure 17. Radio receiver AGC characteristics. 


fied, and selectively subtracted. If the angular veloc¬ 
ity of the wing in yaw or pitch was to be controlled, 
the difference between the appropriate d-c voltages 
would be applied to the differentiating and control 
amplifier exactly as in the case of the photoelectric 
homing control (see Section 4.4.2). If the angle of 
wing incidence was to be controlled, a feedback volt¬ 
age from a potentiometer mounted on wing actuators 
was applied to the d-c amplifier stage ahead of the 
thyratron motor control. 

The audio tones were 300, 475, 755, and 1,195 c. 
With all tones on, as in the differential method just 
described, beat notes are produced between the mod¬ 
ulating tones and between their beats. The beat be¬ 
tween tone 2 and tone 3, for example, is 280 c, which 
beats with tone 1 at 20 c. Thus there was a tendency 
for the wing to hunt at low frequencies. 

The extraordinarily flat AGC characteristic (Fig¬ 
ure 17) of the MIT-Harvey receiver permitted the 
use of a simpler system in the second version. Again 


four tones were applied but never more than two 
simultaneously. The missile responded to the per¬ 
centage modulation of each tone—one each for up, 
down, right, or left. The tones in the up-down regime 
were selected so as not to produce objectionable beats 
with those in the left-right regime. Since only two 
tones were on simultaneously, only one first-order 
beat could be produced. 

The experimental program with the model range 
described at the beginning of this section showed 
that, provided a computer was used, simple on-off 
control of the wing-incidence angle produced accu¬ 
racy equal to any other mode tried. The final version 
of the control system, therefore, abandoned the thy¬ 
ratron motor control, replacing it with heavy-duty 
relays, themselves controlled by the thvratrons. 


4 7 TESTS OF ROC 00-1000 

4 71 Visual Guiding 17 

Some twenty missiles were expended in an attempt 
to attain high-accuracy direct-sight control. Even 
with the high degree of maneuverability in Roc as 
compared with Razon, there can be little hope of suc¬ 
cessful precision bombardment by this means. The 
intervention of a computing sight is required, and the 
parameters of its computation, involving not only 
the position and velocity of the aircraft with respect 
to the target at the instant of release but also their 
whole regime until the instant of impact, are com¬ 
plicated. 

The success attained with Razon through applying 
rigid limitations on the path of the aircraft after re¬ 
lease made continuance of the visually guided Roc 
program unprofitable. 

The extension of the bombing run over the target 
speaks strongly, especially to crews of bombardment 
aircraft, against the use of Razon or of Roc with the 
proposed visual guiding. Crab produced a good first 
approximation to the problem of parallax. It seems 
clear that a more searching study of the bombsight 
might yield a solution to the problem of evasive ac¬ 
tion. Such a program, unprofitable under pressure of 
war, might well be indicated in peacetime. 


Television Guiding 18 

Ten missiles equipped with television were dropped 
before the termination of hostilities closed the ex¬ 
perimental portion of the activities of NDRC. None 














TESTS OF ROC 00-1000 


89 



16 ------------------ 

0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34 36 

SECONDS AFTER RELEASE 

Figure 18. Pitch oscillations in drop T-2, pistol-grip aiming. 


of these was controlled by the computer regime dis¬ 
cussed in Section 4.6.2. In four of the drops the tele¬ 
vision picture was adequate for guidance. In one of 
the remaining six the failure of the television was 
probably due to faulty tuning and is, therefore, hard¬ 
ly to be charged against the television equipment. 

In the drops where the television operated satis¬ 
factorily, the misses were 69, 266, 119, and 133 ft. 
In the best of these drops, an aiming device was used 
consisting of a gimbal-mounted pistol grip. A mirror 
driven by this grip projected a luminous circle ap¬ 
proximately in. in diameter on the television 
screen. The controlling bombardier encircled the im¬ 
age of the target with this ring. Potentiometers on 
the pistol grip controlled the radio transmitter. The 
differentiation in the missile was eliminated, so that 
the wing followed the control law 

8 = k\ 

where X is the pistol-grip displacement. 

The others were dropped without the aid of any 
lomputer. The second and third missiles, with impact 
errors of 266 and 119 ft respectively, were controlled 
from the airplane with a conventional control stick. 
The fourth was controlled by two men from the 
ground, using knob potentiometers. 


These scores can doubtless be improved with prac¬ 
tice, and plans are established to continue the pro¬ 
gram in the Air Technical Service Command. The 
problem of guiding a television bomb, is, however, 
quite complex. Even with a stationary target and an 
approach angle near the vertical to eliminate fore¬ 
shortening, skill is required comparable with that re¬ 
quired for precise dead-stick landings through un¬ 
known wind strata. 

There can be no doubt that the ability of the mis¬ 
sile to fly with the television axis tangent to the flight 
path is vital. Figure 18 shows the apparent oscilla¬ 
tions in pitch of the target in one of the tests. These 
oscillations extend over about 7 degrees of the field 
of view (20 degrees total) and are typical of most of 
the drops. Without a reconstruction of the trajectory 
it is impossible to be certain whether these oscilla¬ 
tions of the target image represent hunting of the 
center of gravity of the missile. Their amplitude and 
frequency, however, would indicate a curvature in 
the trajectory hardly to be attained by Roc. It is 
much more likely that they reflect the failure of the 
missile to fly with zero angle of attack under transient 
conditions. 

Chapter 1 describes similar difficulties which were 
encountered in the glide bombs of the Washington 
























































90 


ROC AND ITS CONTROLS 


Project. The problem was solved there by adding small 
correcting elevators, solenoid-operated by a rate-of- 
pitch gyro. Such a corrective device is probably ap¬ 
plicable to Roc. The hasty application of this pallia¬ 
tive, however, is obviously not to be recommended 
in time of peace. Rather, what is needed is a search¬ 
ing examination of the transient behavior of the 
whole servo-loop system. Techniques are available 
for such study of all components except the airframe. 
The extrapolation of steady-state data from the wind 
tunnel to the transient phase seems to have proved 
unsound. 


Although this project was incomplete at the end of 
the Division’s developmental activities, it has brought 
out the importance of the basic principles which were 
postulated for it at the outset. It has also shown their 
difficulty of attainment and has discovered a need 
for new aerodynamic study to fill a deficiency in 
knowledge in this field. As the speeds of missiles are 
extended through and beyond the sonic range, this 
exploration of the transient phase of their perform¬ 
ance is likely to become increasingly important. A 
prompt and careful attack on this problem cannot be 
urged too strongly. 




PART II 

COMPONENTS AND ACCESSORY ACTIVITIES 




Chapter 5 

TELEVISION 


si INTRODUCTION 

T he same urge that prompted the Japanese to 
inaugurate their program of suicide missiles 
prompted effort in the United States to undertake 
the development of guided missiles equipped with 
television. Under the Japanese type of thought, it ap¬ 
peared simple and appropriate to provide a missile 
with a trained pilot and to expend him against a ma¬ 
jor target. Under American thinking it appeared that 
the same objective could be obtained by removing 
the pilot to a relatively safe location providing him in 
the missile with television to see the terrain under at¬ 
tack and with a radio link to execute his commands. 

Under each culture it was realized that such mis¬ 
siles are expensive, but as the Japanese were prepared 
to expend trained personnel against major targets, 
so it seemed appropriate for our Services to expend 
major logistic effort against targets of crucial im¬ 
portance. The solution of the problem was less simple 
than either nation realized. The efforts of the suicide 
pilots were not uniformly crowned with success, and 
for exactly similar reasons (see Chapter 4) efforts at 
remote control of missiles with the aid of television 
had not been successful as the war closed. The prob¬ 
lem of flying a missile by remote control is not easily 
solved. 

An attack with such a missile can logically be di¬ 
vided into three phases. 

There is a phase of navigation during which the 
missile must be directed to the general vicinity of the 
target. If the missile is of long range, an assault 
drone, for example, then the information conveyed to 
the directing pilot by means of the television link 
must be adequate to give him the same degree of 
recognition of the terrain which is required in contact 
flying. Such information is not always available, even 
with direct vision. The added restriction due to loss 
of resolution and contrast in television, together with 
its limited field of view, further increases the prob¬ 
lem. Probably television is inadequate for this phase 
of the attack, and other methods, such as maintain¬ 
ing visual contact between the drone and its shepherd 
or tracking by radar, would have to be used. 

The second phase of the attack begins when the 
missile is in the general vicinity of the target, which 


has to be distinguished from its surroundings. This 
phase of the attack presents a problem with missiles 
of moderate range (such as glide bombs), as well as 
with long-range missiles (such as an assault drone). 
During this phase there must be no compromise with 
picture quality. The target must be recognized early; 
otherwise, course errors may develop which will be so 
great by the time positive identification takes place 
that subsequent correction of them is impossible. 

In the third and final phase of the attack the pur¬ 
pose of the television is to give the controlling pilot 
continuous knowledge of the course of the missile and 
the error which is likely to take place, so that he can 
continuously correct the error by the application of 
appropriate controls. This phase of the attack pre¬ 
sents a problem in a missile such as Roc (see Chapter 
4). Navigation to the vicinity of the target is made by 
the bombardier before release. For an attack of this 
nature, demands on television are at a minimum; the 
bombardier having identified the target, he is much 
more likely to recognize it when it appears on the 
screen than in the case when its appearance on the 
screen represents his first view of it. Thereafter, his 
problem is simply to see that the target is properly 
positioned on the screen to insure an accurate hit. 
The problem of properly positioning the image of the 
target proves to be a difficult one. Even with missiles 
such as Roc, which fly with zero angle of attack, an 
attempt to keep the target centered on the receiver 
screen leads to failure; some type of lead computer 
is required. For missiles with an angle of attack which 
varies with the amount of course correction required, 
the problem is even more difficult. 

The Division concentrated its efforts in the field of 
television-guided missiles on the high-angle dirigible 
bomb and on Roc, after early experiments with Robin 
indicated but little promise of successful use of a 
television-guided bomb during phase 2 (see above) 
of an attack. 

In addition to developing television for its own 
missiles, the Division cooperated with the AAF in the 
application of television to their glide bomb GB-4 1 
and with both the Army and Navy in the application 
of television to other military uses. Because these ac¬ 
tivities were not a primary divisional responsibility, 
they are not reported here. 


93 


94 


TELEVISION 


52 GENERAL CONCLUSIONS 

The experiments indicated that television was gen¬ 
erally practicable for use as a guiding agency for a 
wide variety of vehicles, within limitations which are 
discussed in this chapter. 

The general reliability of the television system for 
all conditions under which it is to be used is impera¬ 
tive. The great complexity of television apparatus 
and the compactness necessary for military applica¬ 
tions make this requirement difficult to meet, but the 
indications were that after enough effort had been 
put on a given television system it could be made 
reasonably reliable in production models. This effort 
had to be exerted again for each new system de¬ 
veloped. 

It was found essential to the successful use of 
television on guided missiles that the role which 
it is to play be carefully considered in planning the 
operational tactics with the missile. In particular, it 
is easy to underestimate the picture quality which is 
needed to carry out a given steering operation, es¬ 
pecially in the early stages of the maneuver. Further¬ 
more, television viewing is extremely susceptible to 
weather conditions. 

General navigation with television was not found 
to be feasible. The field of view is too narrow, and a 
good quality picture (even with a high-grade system) 
is not enough to permit general recognition of terrain 
over a wide area. An ability to move the camera 
about under control of the operator was found help¬ 
ful in some applications. The extent to which limited 
navigation is practicable was, however, not studied 
extensively by the Division. 

There is a second stage in the maneuver during 
which it is necessary to find the previously unseen 
target. It is to aid guiding at this stage that the best 
quality of picture consistent with the other require¬ 
ments of the television system is needed. The better 
the picture, the less conspicuous a target and sur¬ 
rounding terrain it will be possible to identify. Even 
conspicuous targets and terrain are not too easy to 
recognize at this stage with television equipment of 
weight and compass permitting its carriage in a 
guided missile. 

The third stage consists in guiding the missile to 
the target. Here the requirement on quality is not 
particularly severe, and the principal requirement is 
adequate functioning of the system. 

It was found that after a good television picture is 
obtained it is still necessary to study carefully how 


to steer the missile by it. This difficulty is usually 
underestimated. Airborne missiles are in general not 
so easy to guide as automobiles for example. 

For accurate steering it is necessary that the pic¬ 
ture should show either the instantaneous heading or 
the ultimate destination of the vehicle on the terrain 
with no further steering changes. The best method 
for insuring this characteristic appeared to be to in¬ 
corporate it in the design of the vehicle, i.e., to make 
the missile axis continuously tangent to the line of 
flight. Of the other methods studied, it could only 
be said that it is not enough to compensate for yaw 
without taking account of large transient oscillations 
in the yaw when the bomb is undergoing steering 
changes. 

Indications were that severe wind would influence 
marksmanship, and that automatic computation for 
this and target motion would probably be desirable. 
Only a beginning was made in the study of a com¬ 
puter to correct for these factors. 

Skill and practice in the art of guiding from the 
television picture appeared to be necessary, in the 
stages of development of the apparatus used, to se¬ 
cure accurate hits. A trainer was developed (see 
Chapters 4 and 10) to assist in this. The radio link 
which carries the television must be carefully de¬ 
signed, with the following considerations in mind: 

1. It is easier to build equipment for the longer 
wavelengths, but these give more trouble with an¬ 
tenna directivity. Enough work was done with sys¬ 
tems of 1,000 to 2,000 me to indicate that these are 
generally practical, but more work is necessary to ob¬ 
tain a wholly adequate link from them. 

2. The power output of the system must be large 
enough to obtain good signals at the receiver. The 
problem of jamming was not considered extensively, 
but some immunity to it was developed in the course 
of the work. 

3. The directivity of the antennas (transmitting 
antenna on the missile and receiving antenna at the 
control point) must be broad enough to permit ma¬ 
neuvering both of the vehicle and of the control point 
(when on a plane) to permit continuous transmission 
without loss of signal in extreme angular positions. 
The directivity must, however, be narrow enough to 
eliminate paths involving ground reflection. In par¬ 
ticular, in a fast-moving, medium- or high-angle mis¬ 
sile the forward radiation must be kept down to a 
very low figure. A sufficiently low forward radiation 
was not obtained in the work of the Division with 
high-angle dirigible bombs and with Roc. 



GENERAL CONCLUSIONS 


95 



Figure 1. Block diagram of one video transmission channel. 


4. As between amplitude and frequency modula¬ 
tion, the former is not easily practicable in the higher 
frequency range, i.e., 2,000 me. Frequency modula¬ 
tion will usually give the better results when spurious 
multipaths are kept sufficiently down and the signal 


is strong. The amplitude modulation system, how¬ 
ever, is generally simpler and lighter and does not 
fail so abruptly under adverse conditions. 

The different systems of television tested were 
found, with certain exceptions, to be not too far dif- 


































































































96 


TELEVISION 


ferent. The Iconoscope was found to be a good all- 
around pickup tube. The Image Dissector pickup 
tube is less sensitive than the others but permits some¬ 
what simpler circuits. The Image Orthicon is ex- 
tremeR sensitive but more complicated. 

Infrared sensitivity in the pickup tube gives only 
a slight advantage in seeing through haze, and this 
advantage vanishes completely as the haze becomes 
thicker and turns into fog. This refers to radiation 
about 1 micron in wavelength. There appears to be 
no advantage in utilizing immediately longer wave¬ 
lengths, because of water-vapor absorption. No 
studies were made with television using much longer 
wavelengths. 

53 THE STATE OF TELEVISION AT THE 
BEGINNING OF THE PROGRAM 

531 Television Equipments 

and Circuit Arrangements 

From its earliest experimental stages television has 
captured the American imagination. Even during the 
early days of mechanical scanning, the public refused 
to believe that television was not “just around the 
corner.” Indeed, the feeling was not uncommon that 
television was being viciously withheld from the 
public by a selfishness of the large corporations not 
easily understood or explained. By the outset of the 
war, however, experimental telecasting of pictures 
was regularly taking place in several cities. New tech¬ 
niques in UHF transmission, the advent of the stor¬ 
age camera tube, and the cathode-ray presentation 
tube were the major factors which had made such 
experimental telecasting possible. In spite of these 
developments, however, there was no commercial 
broadcasting of television programs in the same sense 
that sound programs had been on the air fifteen 
years earlier. 

The equipment for broadcasting television pro¬ 
grams 2 was not simple. Figure 1 shows a typical block 
diagram of the equipment required for such experi¬ 
mental broadcasting as was then practiced. This 
equipment was installed at the Radio City studio of 
the National Broadcasting Company and at a trans¬ 
mitting station in the Empire State Building. In ag¬ 
gregate the weight was several tons. The power re¬ 
quirement was in excess of 100 kilowatts, and during 
all broadcasts the entire equipment was under the 
continuous supervision of more than a score of com¬ 
petent technicians. 


As the cuts indicate, the results with this equip¬ 
ment were wholly satisfactory. Figure 2 shows a test 
chart to be transmitted from the television station 
and a photograph of the picture received on the view¬ 
ing screen at a location approximately 45 miles awaj". 
The equipment was adjusted to have a scanning fre¬ 
quency of 525 lines interlaced and 30 frames per 
second. The test chart was illuminated at a level of 
800 footcandles and the camera was equipped with 
a lens of / 4.5. Transmission was on a carrier of 
51-25 me, employing vestigial sideband transmission 
in a channel 6 me wide. 

The equipment indicated in Figure 1 is typical of 
several prewar television installations in the United 
States, and the results shown in Figure 2 are typical 
of prewar television transmission generally. The 
problem of applying television to guided missiles was 
to compress the equipment shown in Figure 1 to the 
compass permitted by the missile geometry, or to 
develop alternative equipment of smaller compass. 
Of equal importance and even more acute was the 
problem of improving the reliability of the equip¬ 
ment so that its satisfactory operation when unat¬ 
tended would be assured. Thus equipment of several 
tons had to be reduced to not over 150 lb, with in¬ 
creased reliability and no serious compromise with 
quality of picture transmission. 

Quality of picture transmission is not readily de¬ 
finable. It consists, essentially, of perfect synchro¬ 
nism in scanning between the camera and the re¬ 
ceiver. Loss of synchronism in horizontal scanning 
results in a sidewise drift of the picture; loss of syn¬ 
chronism in vertical scanning can cause the received 
picture to drift vertically. Phase displacement of the 
scanning will cause improper framing of the picture. 
Simultaneous loss of synchronism in both the hori¬ 
zontal and the vertical sense will result in destruc¬ 
tion of the picture, “tearing.” Continuous and per¬ 
fect synchronization, then, is necessary to receive any 
picture at all. A picture having been received, its 
quality is perhaps best defined by four properties: 
(1) resolution, (2) contrast, (3) brightness, and 
(4) flicker. 

1. Overall resolution is measured by the number of 
horizontal lines which can be distinguished in the re¬ 
ceived scene. This value corresponds to the screen 
fineness of the photoengraver. It is not expressed in 
lines per lineal dimension in the case of television, 
however, on account of the varying degree to which 
the received picture may be magnified. The resolu¬ 
tion obtained is dependent upon many factors, a few 



THE STATE OF TELEVISION AT THE BEGINNING OF THE PROGRAM 


97 


of which are: the resolving power of the optical sys¬ 
tem of the camera, a quality of the photosensitive 
element akin to photographic grain, the density and 
distribution of secondary electrons in the pickup 
tube, the width of band-pass in the radio sections of 
the transmitter and receiver, and such characteris¬ 
tics of the presentation tube as fineness of focus of 
the electron beam and the qualities of the phosphor 
on its screen. Figure 2 shows overall resolution in the 



Figure 2. Resolution chart for testing television trans¬ 
mission (top); photograph of received picture (bottom). 


received picture of 184 lines. This quality was never 
equaled in the missile-borne television equipment 
developed by the Division. About the minimum reso¬ 
lution reasonably tolerable is 120 lines, and with such 
scanning most of the picture’s detail is lost; 240 lines 
represent about the lowest limit of satisfactory reso¬ 
lution when any detail is required; 480 lines give a 
resolution roughly comparable to that obtained from 
the usual 16-mm amateur motion picture. The fore¬ 


going criteria are due to Engstrom ; 3 they have not 
been standardized by the industry. 

2. The maximum contrast obtainable is determined 
by the characteristics of the presentation tube. With¬ 
in the limits set by the presentation tube the con¬ 
trast is determined by the video gain between the 



Figure 3. Change in brightness threshold with scene 
brightness. 


camera and the modulatot of the transmitter and be¬ 
tween the second detector of the receiver and the 
viewing tube. Suppose an image is to be transmitted 
of a perfectly black and perfectly white checker¬ 
board. The portions of the photosensitive surface on 
which the black squares are projected will receive no 
light, and therefore develop no signal to the video 
amplifier. The white squares will excite the photo¬ 
cathode and develop a certain voltage level. With a 



-TWO FRAME CYCLES-► 

Figure 4. Screen illumination for intermittent pictures: 
motion pictures of 24 frames per second (top); televi¬ 
sion at 24 frames interlaced (bottom). 

very low gain the difference between the blackest 
black and the whitest white, one definition of con¬ 
trast, is very small; with a considerable video gain 
the contrast can be considerably increased. Contrast 
is also affected by the received-picture brightness. 
Figure 3 indicates a characteristic of the typical eye 










































98 


TELEVISION 


in resolving halftones. It shows that unless the high¬ 
light brightness of the received scene is adequate, loss 
of contrast will result, irrespective of video gain. The 
quantitative incorporation of this characteristic into 
design awaits further study. 

3. Brightness of the received picture depends upon 
the characteristics of the phosphor screen, the im¬ 
pressed voltage, and a number of other characteris¬ 
tics of the presentation tube and its associated cir¬ 
cuits. This dependence is, of course, based on the as¬ 
sumption that the received signal is of sufficient 
strength to drive the tube to its full capability. Pre¬ 
war tubes had a brightness level too low to be ade¬ 
quate for phase 2 of a guided-missile attack (see 
Section 5.1). The work on cathode-ray oscillograph 
tubes of increased brightness for radar applications 
was of great benefit to the television program. 

4. Flicker is also a function of brightness. Direct 
transference of motion-picture experience to tele¬ 
vision is unsafe: the 24 frames standard for modern 
motion pictures is inadequate to prevent objection¬ 
able flicker in television. This is partly due to the 
difference in the illumination characteristic of the 
viewed presentation: in motion pictures the screen is 
fully illuminated during the entire exposure; with 
television there is an exponential degradation of the 
highlights, starting immediately after a picture ele¬ 
ment has been scanned (Figure 4). 

High-quality television is obtained by 60 repeti¬ 
tions per second of the scene on the presentation 
screen. The use of 500 scanning lines and 60 frames, 
however, requires a band-pass of about 20.4 me. The 
use of interlacing—scanning alternate views of the 
scene with alternate numbered scanning lines—cuts 
the required band in half while preserving the resolu¬ 
tion and freedom from flicker. 

5 3 2 Camera Tubes 

Two types of photosensitive tubes were used for 
the pickup of television scenes for prewar broadcast. 
The first type was the nonstorage type. This repre¬ 
sented a development from the earliest phases of tele¬ 
vision, when mechanical scanning was employed and 
elements of the scene were successively allowed to 
register on the cathode of a photoelectric cell. The 
output current was amplified and modulated the car¬ 
rier wave. 

In the more modern version (the Image Dissector) 
the photocathode is enlarged so that the entire scene 
is focused on it. A potential applied between the 


cathode and an anode produces an electron current 
emanating from the cathode. In space cross section 
this current corresponds in intensity to the level of 
illumination on each element of the photocathode. 
An axial magnetic field focuses the electron image. 
Transverse magnetic fields deflect it horizontally at 
the line-scanning frequency and vertically at the 
frame-scanning frequency. A small aperture (0.02x 
0.02 in.) permits small portions of the space cur¬ 
rent to enter an electron multiplier, which develops 
a signal consisting of a level of current for each 
0.0004-sq-in. element of the electron image. Magni¬ 
fication by electron optics between the photocathode 
and the aperture permits a number of scanning lines 
greater than the quotient of the vertical dimension 
of the optical image or the photocathode by 0.02 in. 
A large number of stages in the electron multiplier 
yields a usable sensitivity. 

The other general class of tubes was known as 
storage tubes. 4 In these devices the photosensitive 
surface consisted of a mosaic cathode. A scene focused 
under such a cathode caused charges to be built up 
on each picture element by the emission of photo¬ 
electrons. An electron gun in the envelope scanned 
the photosurface horizontally and vertically under 
the influence of magnetic fields. This electron beam 
knocked secondary electrons out of the mosaic in 
varying amounts, depending upon the number of 
photoelectrons already emitted. The mosaic was 
backed by, but insulated from, a conducting plate. 
Thus the changing charge on the mosaic-backplate 
condenser formed a signal which was used to modu¬ 
late the carrier. The tube, the Iconoscope, is charac¬ 
terized by high resolution and good contrast at scene- 
brightness levels much lower than were permissible 
with the Image Dissector or with mechanical scan¬ 
ning. It was in production at the beginning of World 
War II in considerable quantity. 

The Iconoscope has an unfortunate characteristic 
known as shading—the appearance of dark areas in 
the received picture which sometimes spread to cover 
the entire viewing screen. The cause of the phenom¬ 
enon is the formation of clouds of secondary elec¬ 
trons released by the mosaic. The mechanism is not 
wholly understood nor has a completely satisfactory 
cure been devised. 

To correct these and other difficulties, a revised 
type of tube was under development at the outset of 
World War II; this tube was known as the Orthicon. 
In this structure a decelerating anode was located 
adjacent to the photosensitive cathode so that the 



THE STATE OF TELEVISION AT THE BEGINNING OF THE PROGRAM 


99 


electrons of the scanning beam impinged on the 
photosensitive surface at substantially zero velocity. 
Thus no secondary electrons were emitted. Further, 
the “stored” photoelectrons emitted by the photo¬ 
mosaic during intervals between the passages of the 
scanning beam were more effectively used, so that 
the sensitivity of the Orthicon was about four times 
that of the Iconoscope. 

The sensitivity of a television camera is conven¬ 
iently expressed by the average intrinsic brightness 
of a received scene which will result in an acceptable 
signal for modulating the carrier. The following tabu¬ 
lation gives the relative thresholds of operation of the 
foregoing tubes with lenses of equal / number. Equal 
/ numbers give an equivalent basis for comparison of 
absolute sensitivities without regard to the logistic 
effort involved in their procurement. Certain tubes 
require a long-focus-length lens which for the same/ 
number is more difficult to procure. 


Image Dissector with electron 
multiplier 

Scene brightness 
required 6,6 
(candles per sq ft) 

10 3 

Iconoscope 

1.0 

Orthicon 

0.25 

Scene illuminated by bright sunlight 

10 4 

Moonlight 

5X10- 3 

Threshold of human eye 

10~ 7 


5 3 3 Transmission Links 

All television broadcasting at the outset of World 
War II followed the conventional UHF AM prac¬ 
tice. Specifically the frequency channels assigned 
were in the 40- to 60-mc band. A channel of 6-mc 
width was assigned to each transmitter. The original 
practice of double sidebands of 2 to 2.5 me had been 
superseded by a single sideband of 4 me in the in¬ 
terests of better resolution. 

In military applications of television, however, it 
was recognized that not all the refinements normally 
invoked in experimental television having commer¬ 
cial application in mind could be employed. Specifi¬ 
cally, interlaced scanning was eliminated, with con¬ 
sequent increase in flicker. A concession in picture 
detail was made by the reduction of the number of 
scanning lines from about 500 to 350. The frame rep¬ 
etition rate of 30 per second was raised to 40 in par¬ 
tial compensation. The resulting required bandwidth 
was approximately 5 me. 


The usual transmission techniques available in ex¬ 
perimental television could not, in general, be applied 
directly to military uses. Some of the work on radio 
relay was, however, applicable. Initial allocations of 
frequency were in the vicinity of 100 me. This was 
soon changed to 300 me and by the end of 1943 the 
only frequencies available for television work were in 
the vicinity of 800 and 1,800 me. 


5 3 4 Receivers 

The typical television receiver in 1939 consisted of 
a superheterodyne circuit in the 40- to 60-mc band 
with the necessary bandwidth designed into the i-f 
amplifier. Intermediate frequencies were rectified in 
a second detector and amplified in a video amplifier; 
the resulting output was applied to the control grid 
of a cathode-ray oscilloscope. A pulse transmitted 
with the video signal at the beginning of each scan¬ 
ning line triggered a local oscillator, whose output 
deflected the electron beam horizontally across the 
screen; at the end of the line-scanning period a blank¬ 
ing pulse returned the line to its original position. 
At the beginning of each frame another pulse, trans¬ 
mitted with the video signal, started the electron 
beam downward across the screen, but at a much 
slower rate; at the end of the frame a blanking pulse 
returned the beam to its position at the upper left- 
hand corner of the picture. A series of pulses, then, de¬ 
flected the electron beam at more or less uniform ve¬ 
locity across the fluorescent screen of the tube so that 
the electron screen “painted” the tube in narrow 
horizontal lines closely placed together. In the ab¬ 
sence of any video signal, the pattern received con¬ 
sisted simply of the horizontal lines. With very close 
scanning, these lines merged so that the received pat¬ 
tern was a clear illuminated rectangle. 

Introduction of a video signal produced an alter¬ 
nating component in the velocity of the cathode ray 
so that the brightness varied from point to point 
along the line and among the lines. In this manner 
the received picture was produced. 

The presentation of any recognizable picture on 
the viewing tube was always dependent, then, upon 
the proper reception of the synchronizing and blank¬ 
ing pulses. Noise in the circuit from static or other 
sources readily produced pulses of a magnitude com¬ 
parable with the synchronizing and blanking pulses. 
Under these circumstances, the picture was de¬ 
stroyed. 



100 


TELEVISION 


54 THE APPROACH OF THE DIVISION 
TO THE PROBLEM 

541 General Overall Engineering 

Even before the formation of the Division, prelim¬ 
inary work had been undertaken by NDRC and by 
the Services toward the development of compact, ex¬ 
pendable television-transmitting equipment for mis¬ 
siles and for other purposes. The first responsibility 
of the Division in regard to television was to coor¬ 
dinate these activities. The compromise of picture 
quality, particularly in resolution, with compactness 
of equipment was in some cases made ill-advisedly, 
and a major initial step of the Division was to reduce 
all television activities to a comparable basis of 350 
lines and 40 frames sequentially scanned. 

A second major overall activity of the Division in 
the television field was to improve the reliability of 
the equipment through the application of sound en¬ 
gineering principles. A great many detailed problems, 
minor individually but frustratingly important in the 
aggregate, were solved. One of significant importance 
was the elimination of crosstalk between the hori¬ 
zontal and vertical sweep circuit in the camera and 
in the receiver. No single activity contributed more 
to the reliability of airborne television equipment 
than the application of proven engineering principles 
to the design of equipment already in advanced de¬ 
velopment. 

5 4 2 Comparison of Camera Equipments 

An attempt was made to rate objectively the sev¬ 
eral pickup tubes available at the beginning of the 
Division’s program and to establish objective rating 
standards for future camera tubes which might be 
developed. This proved to be impractical, however, 
and the comparison of the overall camera equipment 
—optics, pickup tube, scanning equipment, and vi¬ 
deo amplifier—was undertaken. This was a continu¬ 
ing project; 6 the contractor’s report is recommended 
for study to all those interested in the application of 
television to military problems. Sensitivity of various 
equipment was measured and the performance of the 
equipment from threshold to a highlight brightness 
of about 1,000 foot-Lamberts explored. No objective 
measure of contrast similar to the gamma of the pho¬ 
tographer was known. Contrast was rated as good, 
low, or high. Brightness of the received picture and 
flicker were not assessed. 

An attempt at a quantitative evaluation of con¬ 


trast was made by measuring on a cathode-ray oscil¬ 
lograph the video output of the camera when scan¬ 
ning a strip of white blotting paper on a matte back¬ 
ground and also by measuring the variation in signal 
developed by light transmitted through a translucent 
gray scale having ten logarithmic steps from 2 to 0.2. 
In the latter test certain of the conversion equip¬ 
ments show perceptible linearity between the log¬ 
arithmic density of the translucent gray filter and 
the signal strength developed. It appears that further 
work along these lines might reasonably be expected 
to yield an objective contrast rating. 

The importance of using objective, quantitative 
methods of evaluating television performance can 
hardly be overemphasized. 

The several camera equipments were also rated as 
to blocking, shading, and distortion. Spectral re¬ 
sponse was taken by measuring the response at vari¬ 
ous .levels of illumination at a color corresponding to 
5,400 K and with the following filters. 

Wratten 47 Blue 

Wratten 61 Green 

Wratten 29 Red 

Wratten 88A Near-infrared-passing 

Aklo 9780 Infrared-blocking 

The performance of the equipments under me¬ 
chanical vibration at low frequencies was not meas¬ 
ured, but each unit was placed in an acoustic cham¬ 
ber and subjected to very high noise levels from a 
loudspeaker. The noise level in db above 10~ 16 watt 
per sq cm which produced interference was deter¬ 
mined at several audio frequencies from 400 to3,500 c. 

5 4 3 Improvements in Sensitivity 

Two approaches are available to improve the sen¬ 
sitivity of television pickup equipment. The aperture 
of the optical system can be increased, permitting 
more light to fall on the photosensitive material. 
This is normally accompanied by loss of resolution, 
increase in the complexity of the lens system, or 
both. This is even more true in the case of television 
than in the case of photography, since the usual 
photosensitive surface has a broader spectral response 
than most photographic film, so that color correction 
in the lenses must be extended into the near infrared 
region. A second approach to the problem is through 
the increased sensitivity of the pickup tube itself. 

Both these avenues were explored by the Division. 
Lenses of / 3.5 were customary in television work. 
Further extension in this direction appeared unprofit- 



THE APPROACH OF THE DIVISION TO THE PROBLEM 


101 


able, since the geometry of the camera tubes requires 
a lens of a considerably greater focal length than 
would be called for by the diagonal of the projected 
image. Schmidt optics with an equivalent aperture of 
about / 0.7 was applied to each of the pickup tubes 
available. 

A major improvement, therefore, of the camera 
pickup tube itself seemed desirable. Preliminary work 
had been started before the war on the Image Orthi- 
con 5 tube. This tube was a modification of the Orthi- 
con, in which an electron image was created on a glass 
target behind the photocathode. Electrons from the 
photocathode impinged on the target with sufficient 
velocity to cause the emission of secondary electrons. 
These electrons were collected on a fine screen, pro¬ 
ducing a residual positive charge on the target. A 
rise in the target potential was prevented by the ad¬ 
hesion of some of the electrons of the scanning beam. 
The insulating property of the target caused the 
charge distribution, and therefore the capture of elec¬ 
trons from the scanning beam, to correspond to the 
electron image impinging on it. The electron beam 
was thus modulated by subtraction at the target. 
The modulated return beam was collected and ampli¬ 
fied by electron multiplication to produce the signal. 

The target was made extremely thin (0.2 mil) so 
that a moderate leakance existed between the front 
(electron-image) and back (scanned) surfaces. Thus 
the charge resulting from the loss of secondaries was 
neutralized in each scanning cycle without produc¬ 
tion of dangerous voltage stress between the target 
surfaces. 

5 4 4 Television Radio Links 

Transmission at Higher Frequencies 

At the beginning of World War II, television ex¬ 
perimentation for military purposes was at 100 me. 
This channel was very early taken over by other 
communication services, and television was pushed 
to the 300-mc band. As warfare progressed and com¬ 
munication and radar required more and more of the 
spectrum, the only bands finally available for tele¬ 
vision were two—one about 800 me and the other 
about 1,800 me. In addition some work was done on 
television transmission at 1,200 me. 

Multipath Problems 

At the outset, multipath transmission caused con¬ 


siderable difficulty. It resulted in the appearance of 
vertical bars in the received picture, which could only 
be accounted for by beating of the signal transmitted 
over the longer reflected path with that transmitted 
over the direct path. Considerable analysis of the 
video signal was made to ^eek out a source of the 
beats near the portions of the scene where there was 
an abrupt change in the video frequency. The major 
cause, however, proved to lie in a spurious frequency 


DIRECT SIGNAL 



BOMB IMAGE 


/y 

' / v 

Figure 5. Multipath interference. 


modulation of the carrier, resulting from close cou¬ 
pling of the master oscillator with the modulator. To 
save space the usual buffer stage separating the oscil¬ 
lator from the modulator had been eliminated. The 
reaction of the modulator on the oscillator resulted in 
frequency pulling of the order of 100 kc. The restora¬ 
tion of the buffer reduced the spurious FM to the 
order of 10 kc, well below the threshold at which 
multipath difficulties are detectable. 

Even with a fixed carrier frequency, however, mo¬ 
tion between the transmitter and the receiver will in¬ 
troduce a changing frequency through the doppler 





102 


TELEVISION 


effect. With missiles of moderate speed such fre¬ 
quency shifts are small, and the effects on picture 
transmission are negligible. With the high-angle dir¬ 
igible bomb, however, the bars characteristic of 
spurious FM were present. (See Section 5.6.2.) In 
many cases the beating was so pronounced as to cause 
loss of synchronism and tearing of the picture. 



Figure 6. Block III equipment: camera (top), trans¬ 
mitter (center), and receiver (bottom). 


The major difficulty which confronted the Divi¬ 
sion’s investigators was beating between the signal 
from the transmitter in the bomb and the signal from 
its reflected image (Figure 5). This interference pro¬ 
duces bars across the picture whose direction depends 
upon the relative change in length of the two trans¬ 
mission paths. At the velocities of the high-angle 
dirigible bomb and Roc, the bars are horizontal and 
of considerable intensity. 

If an application of television to high-speed mis¬ 
siles is to be considered, this problem, as yet un¬ 
solved, must receive careful attention. Preliminary 
work in this direction was done with higher frequencies 
(1,800 me) in an effort to eliminate downward radia¬ 
tion from the missile and the image of its transmitter. 

Frequency Modulation Vs Amplitude 
Modulation 

The pronounced improvement in reception of au¬ 
dio signals when transmitted by FM as compared 
with AM spoke strongly for the exploration of this 
type of transmission for video signals. The Division’s 
contractors made comparative tests with both types 
of transmission at 300 me, at 775 me, and at 1,200 
me. At 1,800 me the geometry of the tubes in the r-f 
circuits precludes amplitude modulation. 


5-5 SUMMARY OF RESULTS 

5,5,1 Equipment Simplification 

and Improvement 7 >W° 

The elimination of crosstalk between the horizon¬ 
tal and vertical scanning circuits and the video am¬ 
plifier has already been mentioned. It was accom¬ 
plished largely by cleaning up the wiring of the cam¬ 
era, transmitter, and receiver equipments (Figure 6). 
This step, in addition to the elimination of spurious 
FM by the decoupling of the oscillator and modula¬ 
tion in the transmitter, represented a major contribu¬ 
tion to the reliability of television at Block I and 
Block III frequencies (100 me and 300 me). 

The early models of military television equipments 
were very discouraging when flight-tested; one cause 
was the spurious FM already mentioned; the second 
was low percentage modulation. The correction of 
the latter difficulty was complicated by the fact that 
no ready means was available at the outset of the 
program of measuring percentage modulation at the 
broad band of modulating frequencies occurring at 














SUMMARY OF RESULTS 


103 


the transmitter. The accepted practice was to apply 
sine-wave modulation and to assume that the per¬ 
centage modulation remained unchanged with the 
change in waveform of video modulation. A percent¬ 
age-modulation meter was improved from which the 
approximate percentage modulation could be scaled 
off. 

The application of appropriate engineering to the 
production design of the Block III resulted in the 
elimination of six of the nine operating control ad¬ 
justments, without loss of ability to select any one 
of ten channels with a single equipment. The same 
simplification increased the output power by 39 per 
cent from 18 to 25 watts. The corresponding increase 
in range of 18 per cent is perhaps not importafit; 
however, the improvement in signal-to-noise ratio at 
ranges less than the limiting range is very important 
in reducing the effects of interference. 

The r-f end of the receiver was redesigned, with a 
much improved head-end tuner. The new tuner per¬ 
mitted the use of an r-f stage with an improvement 
in the ratio of the signal to its image of from 3.5 to 
100. The overall result was a 14-db improvement in 
signal-to-noise ratio. 

The receiver scanning oscillators were stabilized, 
and an improved AGC circuit was developed. The 
original cathode-coupled multivibrators were replaced 
by sine-wave oscillators capable of automatic syn¬ 
chronization with the received synchronizing pulses 
by AFC circuits. With the added AFC and the stabi¬ 
lized oscillators there was no tendency for the received 
picture to tear, even with loss of synchronizing pulses 
for several frames. New low-impedance AGC circuits 
were designed with low time constants in comparison 
to a line duration, in contrast with the high- 
impedance AGC circuits which they replaced. They 
resulted in successful reception of signals through 
pulse interference, even if the pulse had an amplitude 
1,000 times the desired signal, provided the interfer¬ 
ing pulse rate was not too close (within 1 or 2 per 
cent) to the scanning frequencies. The combination 
of low-impedance AGC and the stabilization of the 
scanning oscillators with AFC resulted in maintain¬ 
ing protection against 4,000-to-l pulse interference, 
even with pulse repetition rates within 0.4 per cent 
of scanning frequencies. AFC had not been incor¬ 
porated into production receivers as the war ended; 
stabilized oscillators and fast AGC improvements 
were. 

A further improvement of about 4 to 1 in allow¬ 
able peak noise-to-signal ratio was obtained by the 


use of a keyed low-impedance AGC, which permitted 
signals to pass through the AGC amplifier only dur¬ 
ing the synchronizing pulse. This kept the bias con¬ 
stant during the line duration. At the close of World 
War II, this form of AGC had not been applied to 
military television receivers, its dependence upon 
very precise maintenance of synchronism having de¬ 
ferred its application. The use of the keyed low- 
impedance AGC during periods of poor synchronism 
resulted in an effect similar to motorboating. 

A limiter was applied to the receivers to clip noise 
peaks to a value only slightly higher than the syn¬ 
chronizing peaks. A second limiter, acting on the 
video input to the synchronizing circuits, limited the 
synchronizing-pulse amplitude at strong signal levels, 
as well as limiting pulse noise to amplitudes equal to 
the S 3 mchronizing pulses with weak signals. This im¬ 
provement, not available until the close of the pro¬ 
gram, further reduced the vulnerability of the radio¬ 
transmission link to pulse noise near the scanning 
frequency. 

The Iconoscope camera tube is limited in latitude 
(the range in scene brightness over which fidelity of 
response is maintained) on account of the clouds of 
secondary electrons already discussed. To accom¬ 
modate scenes which change over a broad range of 
brightness, a lamp was placed behind the mosaic. 

The provision of suitably insulated windings for 
the horizontal output transformer solved the problem 
of a suitable supply for the Iconoscope cathode heater. 
In addition to reducing the high-voltage cabling to a 
minimum, it permitted a simplification in the hori¬ 
zontal deflection output circuit. 

A “clamp” circuit was developed which sets or 
clamps the bias of the output video stage at the be¬ 
ginning of each line and then opens the bias supply 
circuit for the duration of the line. This results in re¬ 
moving all microphonic disturbances in the preceding 
video stages and further permits designing the video 
amplifier so that it is flat only down to line frequency. 
This eliminates most of the electrolytic filter capaci¬ 
tors while maintaining satisfactory reproduction of 
scenes that have frequency components below line 
frequency, e.g., scenes including the horizon. 

The limitations of space in the Block III camera 
prevented its use. In its place an RC network was 
inserted in the grid circuit of the first video stage. 
The time constant of this network, the leveling cir¬ 
cuit, was adjusted to give approximately the same 
overall effect as the clamp circuit. Its efficacy was 
rather less, however, and the clamp circuit was used 



104 


TELEVISION 


on the compact equipment, the Orthicon, and the 
Image Orthicon equipment. 

The availability of 6AG5 miniature pentodes at 
the later stages of the program permitted the design 
of a much more compact video amplifier. Further sav¬ 
ing in space was accomplished by the elimination of 
peaking coils through the use of small capacitors to 
by-pass the cathode resistors. In spite of the loss of 
gain per stage due to degeneration at low frequencies, 
the overall gain of the amplifier strip was preserved 
at sufficient saving in space to permit the reinsertion 
of a buffer amplifier without increasing the overall 
size of the equipment. 

5 5 2 Comparison of Camera Equipments 6 

The original plan under this project was to make 
an objective comparison of television pickup tubes. 
The intimate association, however, of the pickup 
tubes with the remaining portions of the camera 


equipment prevented this objective from being real¬ 
ized. Since the comparisons made, therefore, are be¬ 
tween complete camera equipments, they do not lead 
to a conclusion as to which is the best pickup tube 
to choose for the optimum design of associated cam¬ 
eras. 

Seven camera equipments were tested as follows. 


Name 

Block III 
2-in. Orthicon 

Experimental Image Orthicon 
Production Image Orthicon 
Image Orthicon with Schmidt optics 
Vericon 

Image Dissector with multiplier 


Manufacturer 

RCA 

RCA 

RCA 

RCA 

RCA 

Remington Rand 
Farnsworth 


The equipments under test were assembled in a 
single laboratory where they could be compared un¬ 
der controlled conditions (see Figure 7). Lighting of 
adjustable level and color was provided. The illumi¬ 
nation was continuously adjustable from 0.02 to 
2,000 footcandles. The corresponding intrinsic bright- 



Figure 7. Conversion units set up for comparative measurement of resolution and sensitivity. 
































SUMMARY OF RESULTS 


105 


Table 1 

Block III 

Type of camera tested Iconoscope 

Vericon 

Image 

Dissector 

2-in. 

Orthicon 

Experi¬ 

mental 

Image 

Orthicon 

Image 

Orthicon 

Production with 
Image Schmidt 

Orthicon optics 

Optical system—focal length (cm) 

20.3 

5.0 

23.7 

1.7 

12.0 

8.9 

20 

/ number 

4.5 

1.9 

2.5 

1.2 

27 

3.5 

1.0 

Threshold sensitivity (ft-Lamberts) 

8.2 

18.3 

175 

27 

1.5 

0.5 

0.07 

Optimum highlight brightness (ft-Lamberts) 

280 

210 

420 

600 

12 

8.0 

0.7 

Resolution at threshold 








Area A 

130 

160 

180 

100 


100 


Area B 

160 

140 

150 

100 


150 


Area C 

200 

120 

150 

200 


170 


Area D 

200 

140 

190 

130 


120 


Area E 

200 

200 

200 

200 

200 

200 

100 

Resolution at optimum brightness 








Area A 

260 

290 

200 

160 

300 

200 

120 

Area B 

300 

290 

170 

200 

300 

190 

150 

Area C 

260 

240 

170 

240 

250 

180 

200 

Area D 

250 

250 

200 

198 

250 

180 

150 

Area E 

300 

300 

240 

250 

300 

260 

200 

Resolution at high brightness level 








(ft-Lamberts) 

580 

580 

420* 

600* 

580 

580 

250 

Area A 

260 

290 

200 

100 

300 

110 

120 

Area B 

300 

290 

170 

200 

300 

200 

150 

Area C 

260 

240 

170 

240 

250 

190 

200 

Area D 

250 

250 

200 

198 

250 

160 

150 

Area E 

300 

300 

240 

250 

300 

210 

120 

Contrast at threshold 

Low 

Low 

Low 

Low 

Low 

Low 

Low 

Contrast at optimum 

Good 

Good 

High* 

Good 

Good 

Good 

Good 

Contrast at high brightness level 

Good 

High- 

to 

Good* 

High 

High 

High 



Good 

Good* 






*Same as optimum. 


KEY TO RESOLUTION AREAS 



ness of the highlights in the scene viewed varied be¬ 
tween 0.005 and 500 candles per sq ft. The spectral 
distribution of the illumination corresponded to a color 
temperature of 5400 K. Color sensitivity was deter¬ 
mined by measuring the response of the camera equip¬ 
ment when viewing scenes consisting of Wratten fil¬ 
ters in combination with Eastman logarithmic step 
gray scales. The entire assembly was covered by a 
filter of Aklo glass to cut out the near infrared. 

Sensitivity and resolution were determined by 
viewing a standard 18x24-in. black-and-white reso¬ 
lution chart at a distance which covered the photo¬ 


sensitive surface, with the exception of a small margin 
as recommended by the manufacturers. The test 



too 200 500 1000 2000 5000 10,000 


FREQUENCY IN C 

Figure 8. Noise level above 10 -16 watt per sq cm at 
which microphonics produce serious interference with 
video signal. 













































106 


TELEVISION 



Figure 9. Comparison of resolution of camera equipments: A, Vericon; B, Dissector; C, 2-in. Orthicon; D, experimental 
Image Orthicon; E, Image Orthicon with Schmidt optics; F, production Image Orthicon; G, Block III Iconoscope. 
































SUMMARY OF RESULTS 


107 


chart permits direct observation of resolution be¬ 
tween 100 and 300 lines. Table 1 gives a comparison 
of the observed sensitivity and resolution of the seven 
equipments. 

Contrast was reported as high, good, or low; a 
quantitative evaluation of contrast similar to gamma , 
well-known to photographers, was not established. 


The vulnerability of the camera equipment to in¬ 
terference from audio vibrations was determined by 
enclosing each equipment in a soundproof cabinet 
with a powerful loudspeaker. Figure 8 shows the 
frequencies and noise levels in decibels above 10~ 16 
watt per sq cm at which, serious interference oc¬ 
curred. 



Figure 10. Comparison of amplitude linearity of camera equipments: A, Vericon; B, Dissector; C, 2-in. Orthicon; D, 
experimental Image Orthicon; E, production Image Orthicon; F, Block III Iconoscope. 



















108 


TELEVISION 


Table 1, prepared from data presented in Columbia 
Broadcasting System’s report, 6 indicates satisfactory 
picture reception when the peak-to-peak signal-to- 
noise ratio is less than unity. This ratio has been rec¬ 
ognized as a generally inadequate criterion unless its 
frequency distribution is specified. Broadly it is found 


that more random noise is tolerable in the upper than 
in the lower stretches of the video frequency band. 

The Columbia investigation did not attempt to 
evaluate the data to reach a conclusion as to the 
best overall pickup tube. The greatly increased sen¬ 
sitivity of the Image Orthicon speaks very strongly in 



Figure 11. Comparison of color response of camera equipments: A, Vericon; B, Dissector; C, 2-in. Orthicon; D, experi¬ 
mental Image Orthicon; E, production Image Orthicon; F, Block III Iconoscope. 















SUMMARY OF RESULTS 


109 




Figure 12. Image Orthicon camera with Schmidt 
optics. 


its favor. The overall fidelity of its output at any 
light level, however, is not equal to that of the Icono¬ 
scope. In addition, (1) it is more sensitive to fluctua¬ 
tions of supply voltage, (2) it loses contrast range 
with wide ranges of illumination, (3) it is more sus¬ 
ceptible to microphonics, and (4) it has a lower signal- 
to-noise ratio. This last element does not appear 
from the work of Columbia because of the limited 
frequency range over which the noise was studied. 
The high sensitivity of the Image Orthicon in the near 
infrared portion of the spectrum, 1.0/x, is most at¬ 
tractive for certain military operations. This sensi¬ 
tivity was partly lost, however, in the miniature ver¬ 
sion of this tube. On the whole it appears that the 
best overall performance is obtained from the Icono¬ 
scope, but that for purposes where extreme sensitiv¬ 
ity is paramount the Image Orthicon is to be pre- 


Figure 13. Bench test setup for sealed-beam Dissector 
with Schmidt optics. 


ferred. There are possibilities, also, that for special 
uses some of the other pickup tubes may be advan¬ 
tageous, although no clear case of such use appeared 
in the work of the Division. Figures 9, 10, and 11 
compare resolution and sensitivity, amplitude linear¬ 
ity, and color response of the conversion units tested. 

5 5 3 Improvements in Sensitivity 

The threshold sensitivity of the Block III camera 
using the Iconoscope with an / 4.5 lens was of about 
8.2 ft-Lamberts. This was typical of the maximum 
sensitivity available at the beginning of the Division’s 
program. Motion-picture cameras and emulsions 
were available for work at illumination levels of ap¬ 
proximately one-tenth this level, and there was con¬ 
siderable pressure from the Services to get equal 


television performance. Two approaches to the prob¬ 
lem were made, one by the increase in aperture of the 
optical system through the use of Schmidt optics, the 


Figure 14. Vericon tube and camera (cover removed). 

other by the improvement of the pickup tube itself. 
Schmidt optics applied to the Image Orthicon 11 (Figure 
12) showed an improvement in threshold sensitivity 


VOLTAGE DIVIDER 




110 


TELEVISION 



*2 GRID**) DYNODE ♦200 V 
GRID 

CATHODE OV 
PHOTOCATHODE-600V 
TARGET SCREEN 0 V 
■"S GRID 0 V 


Figure 15. Image Orthicon tube with focusing coils and deflection yoke. 


as reported in Table 1 of 7 times. The optical aper¬ 
ture for the conventional system was / 3.2, and the 
aperture of the Schmidt system was approximately 
/ 1.0. An increase in sensitivity slightly less than the 
square of the / number may be due to losses in the 
correction plate and Cassegrain mirror. 

Experiments were also made in the application of 
Schmidt optics to the low-velocity Iconoscope 12 and to 
the Image Dissector. 13 Success with the former was 
frustrated by the mosaic’s charging at even moderate 
illumination levels. The spherical mirror for the Im¬ 
age Dissector with Schmidt optics was sealed in the 
envelope of the pickup tube in a sort of sealed-beam 
construction (Figure 13). The program was dogged 
with mechanical difficulties and no definite conclu¬ 
sion was reached. 

The second method of improving sensitivity was 
through improvement of the pickup tubes them¬ 
selves. Development of a small Orthicon-type tube 
and camera (Figure 14) was carried out by Reming¬ 
ton Rand 14 under Contract OEMsr-187. This contrac¬ 
tor’s version of the tube is known as the Vericon. 

The resolution of the Vericon is rather better than 
that of the Orthicon. Its sensitivity as tested by Co¬ 
lumbia (see Table 1) is somewhat higher than the 
Orthicon, even without correcting for the greater in¬ 
cident light received by the Orthicon in the Columbia 
tests (/ 1.9 for the Vericon, compared with / 1.2 for 
the Orthicon). It is rather less subject to microphonics 
at audio frequencies; otherwise it is substantially 
similar to the Orthicon. 


The most useful gain in sensitivity was accom¬ 
plished by adding image multiplication and electron 
multiplication to the Orthicon tubes. Table 1 testifies 
to the success of this development. A further im¬ 
provement beyond that noted in the tabulation is the 
sensitivity of the Image Orthicon to the near infrared. 
With a Wratten 88A infrared filter, the Iconoscope 
produced a fair picture, but with rather low contrast 
and resolution at 1,600 footcandles of illumination. 
With the same filter, the Image Orthicon produced a 
picture of slightly decreased contrast at an illumina¬ 
tion level of 23 footcandles. With a Corning 9780 
filter, the corresponding illumination levels were 785 
footcandles for the Iconoscope and 18 footcandles for 
the Image Orthicon. The Wratten 88A filter is prac¬ 
tically opaque to visible light but transmits freely 
beyond 0.71 /x; the Corning 9780 blocks beyond 0.7 /x 
but passes freely in the visible range. This increase 
in sensitivity to the infrared is particularly important 
in outdoor applications at dawn and dusk, when mili¬ 
tary application of television may be of particular 
importance. 

The Image Orthicon tube (Figure 15) has been fully 
described in the public technical press as well as in 
the contractor’s reports. 5 ’ 7 Some of the difficulties en¬ 
countered in its development are worth mentioning. 
The problem of mounting a 0.2-mil glass target on a 
supporting ring so that it would be free from ripples 
was solved 15 by firing a thin section of a blown bubble 
over a metal ring. At softening temperatures, the 
glass draws into a tight drumhead by virtue of its 









































SUMMARY OF RESULTS 


111 


surface tension. A carefully controlled heating and 
cooling schedule was developed to bring the structure 
down to room temperature while the glass was still 
under tension and flat. 

The target screen consists of fine-mesh wire screen 
with a high percentage of opening area. A method 
was developed which produced screens with 1,000 
meshes per in. and 40 to 80 per cent open area. A 
sheet of glass was precision-ruled with the desired 
mesh and covered with a thin layer of metal. The 
metal was then removed from all except the rulings. 
Additional metal was plated onto the lines and the 
resultant screen stripped off. 

An efficient multiplying electrode was designed 
which would accept electrons from one side, multiply 
them, and allow the secondaries to pass to the follow¬ 
ing stage. A “Venetian blind” structure was evolved 
which was substantially opaque to incident electrons 
but which provided liberal apertures for passage of 
the secondaries to the adjacent stage (Figure 16). 
The efficiency of this structure, 80 to 90 per cent, 
was sufficient to provide adequate gain from three 
stages. 

“The present limiting factor of the tube’s perform¬ 
ance is its low signal-to-noise ratio compared with 
the Iconoscope performance. This low ratio results 
from the low target capacitance.” This conclusion, 
quoted from the contractor’s final report, should be 
reexamined in terms of the noise-signal-frequency 
relations discussed above. In any case it appears that 
the overall efficiency of the tube approaches, as far 
as sensitivity is concerned, the quantum efficiency 
of the photosurface. 

5 5 4 Radio-Transmission Links 

Amplitude vs Frequency Modulation 

In an effort to obtain the same improvement in 
freedom from interference for television that has been 
obtained with audio transmission, experiments were 
made with frequency modulation at each of the car¬ 
rier frequencies with which the Division’s contractors 
were concerned. At all except the highest band (1,800 
me) frequency modulation was inferior to amplitude 
modulation. Amplitude modulation at 1,800 me 
proved impracticable. 

At Block I and Block III frequencies, 16 ’ 17 it appeared 
that with the same bandwidth per channel, FM is 
definitely inferior to AM. This is true even with a 
stationary transmitter and receiver; with relative 


motion between transmitter and receiver the picture 
degradation is even greater. The pictures of Figure 17 
show the results of multipath beating. Where large 
changes in brightness took place, the range of the dis¬ 
criminator was sometimes exceeded. Under such cir¬ 
cumstances the swing of the, signal carried over onto 
the reverse slope of the discriminator characteristic, 
producing a negative image. 



Figure 16. Image Orthicon—miniature in foreground 
and standard size in background. 


The comparison of FM and AM at other frequen¬ 
cies is discussed in connection with the development 
of radio-transmission links at higher frequencies. 

Television at Higher Frequencies 

The continuing increase in communication density 
continually drove television into higher and higher 
ranges of the electromagnetic spectrum. The first 
work on military television was carried out at about 
100 me. Even before the organization of the Division, 
this band had to be abandoned, and a band in the 
vicinity of 300 me was assigned. During 1943 the 
300-mc band was barred to television, and bands 
were made available at 800, 1,200, and 1,800 me. 

Block X. 18 This program, carried out by Philco un¬ 
der Contract OEMsr-1159, covered the development 
of a radio-transmission link at 775 me. Video and syn¬ 
chronizing were provided by a standard Block III 
camera unit. Two transmitters were developed, one 
using frequency modulation, the other amplitude 
modulation (Figure 18). Appropriate receivers were 
developed for each transmission system; the video 
end of the receivers (viewing tube, sweep circuits, 
etc.) were taken from standard Block III receivers. 

AM Transmitter. The video signal from the camera 
unit was amplified through three stages and a driver 








112 


TELEVISION 



Figure 17. Multipath, distortion, and reversal of television pictures in FM transmission at 312 me. A and B show 
multipath interference; C shows some distortion; D is a substantially normal picture; E and F are adjacent frames of a 
16-frame-per-sec motion-picture record, F being partially reversed; G and H show almost complete reversal. 






















SUMMARY OF RESULTS 


113 


(Figure 19). The output of the driver (modulator) 
was impressed on the cathode of the final r-f output 
stage. Synchronizing signals were similarly amplified 



Figure 18. Block X AM transmitter. 


through three stages and a modulator, rectified, and 
impressed on the plate of the r-f output stage. The 
r-f system consisted of a master oscillator, a buffer, 
and a single output stage. Oscillator, buffer, and out¬ 


put tubes were 2C44 tubes in resonant cavities with 
small plate and grid trimmers. Degenerative feed¬ 
back around the oscillator and output provided stable 
wide-band operation. Interconnection between stages 
and to the antenna was by short coaxial lines. The 
antenna consisted of two dipoles driven in phase, 
with ^-wavelength vertical Spacing (Figure 20). Two 
parasitic reflectors, each a quarter-wavelength be¬ 
hind the driver dipoles, assisted in producing an over¬ 
all antenna gain of 4. The output of the AM trans¬ 
mitter was 4 watts. 

AM Receiver. The AM receiver was a superhetero¬ 
dyne with 446-A tubes as the oscillator and mixer 
(Figures 21 and 22). Its six-stage i-f amplifier had a 
band-pass of 10 me, centered at 60 me; the three- 
stage video system had a band-pass of 4 me. Fast 
(0.01-sec time constant) delayed AGC provided sub¬ 
stantially flat output between 150-v and 1,000-v 
input. 

FM Transmitter . The r-f suboscillator of the FM 
transmitter (Figure 23) operated at 260 me. Video 
and synchronizing signals were separately amplified 
and mixed in a cathode-follower stage, which drives 
a reactance tube to swing the oscillator frequency 
through V/% me. The modulated oscillator was am- 


v-1 


V-2 V - 3 


V- 4 



Figure 19. Block diagram of Block X AM transmitter. 




























114 


TELEVISION 


plified through two push-pull, Type 832 amplifier 
stages, which drove a cavity-type frequency tripler 
followed by two cavity-type output stages. The an¬ 
tenna used with this transmitter (Figure 24) was 
identical with that used with the AM transmitter. 
An output of 3.5 watts was attained. 



Figure 20. Block X AM transmitting antenna. 


FM Receiver. The FM receiver (Figures 25 and 26) 
was a unit similar to the AM unit with such differ¬ 
ences in the i-f system as were dictated by the use of 
frequency modulation. The local oscillator and mixer 
were identical; they were followed by a total of seven 
i-f stages, the last two of which were limiters. The 
i-f bandwidth was 10 me. The discriminator em¬ 
ployed two 9006 tubes, each with its tuned circuit 
capacity coupled to the last i-f stage. AGC was rather 
less effective than for the AM receiver, the output 


rising some 80 percent as the input signal increased 
from 150 v to 1,000 v. 

Flight Tests. Figure 27 shows the comparative pic¬ 
ture qualitjr for the two systems, and the following 
tabulation gives their comparative performance. 



AM 

FM 

Transmitter power (average) 

4 watts 

3.5 watts 

Range, substantially free of noise 

5 miles 

2.5 miles 

Limit for reliable synchronization 

10 miles 

5 miles 

Maximum range under optimum 
conditions 

32 miles 

20 miles 


Block XII. 19 The improvement of airborne television 
equipment by raising the carrier frequency to 1,200 
me was considered. It was hoped that this change 
would (1) reduce aerodynamic drag by reduction of 
size of the antenna structure; (2) improve directivity, 
thus improving the power radiated in the desired 
direction; (3) obtain more channels and give greater 
freedom from electrical jamming; (4) result in less 
interference from ignition and pulse sources. Lack of 
suitable tubes to permit the design of a broad-band 
amplifier at this frequency blocked the program. By 
the time tubes were available (see Block X above and 
Block XVIII below), this frequency band was closed 
to television. 

Block XVIII. 20 The Division assigned to the Gen¬ 
eral Electric Company under Contract OEMsr-1172 
the task of developing a transmission link in the 
1,850- to 2,000-mc band. Originally it was planned 
that both AM and FM systems would be developed. 
Initial calculations indicated, however, that even if 
the carrier was crystal-controlled amplitude modula¬ 
tion of the final stage would produce sufficient varia¬ 
tion of the transit angle to give an intolerable amount 
of phase modulation. These preliminary conclusions 
were confirmed by AM experiments using tubes of 
lighthouse construction (ZP464 and L30C). The pre¬ 
liminary work, analytical and experimental, further 
showed that plate amplitude modulation at the os¬ 
cillator was a promising method of obtaining fre¬ 
quency modulation (Figure 28). Specifically, 15-mc 
swing of an 1,850-mc carrier could be obtained with 
20 to 30 per cent amplitude modulation. This range 
of AM was well within the scope of well-designed 
limiter stages. 

The frequency-modulated transmitter consisted of a 
reflex velocity oscillator using a 2K28 lighthouse tube 
in a cavity tunable to the carrier frequency. Standard 
2K28 tubes could not be depended upon to give a 
15-mc swing when amplitude modulated 20 to 30 per 












SUMMARY OF RESULTS 


115 



Figure 21. Block diagram of Block X AM receiver. 


cent; selected tubes, however, would yield at least 
this degree of frequency modulation. 

The video and synchronizing signals were brought 
separately to two 6AB7 pentodes, with individual 
gain control obtained through the bias to the indi¬ 
vidual grids. Joint gain control was provided by volt¬ 
age adjustment of the common grid-bias potential 
source. The plates of the 6AB7’s were paralleled for 



mixing the video and synchronizing signals. A 6AG7 
stage amplified the combined video and synchroniz¬ 
ing signal to 60 volts, peak-to-peak for plate modula¬ 
tion of the oscillator. 

Two stages of power amplification, using ZP579’s 
and interconnected with tunable plumbing, supplied 
the antenna. As originally laid out, the transmitter 


VIDEO 

IN 


SYNC 

IN 



Figure 23. Block diagram of Block X FM transmitter. 


supplied 5 watts to the antenna; revisions in the cav¬ 
ity and plumbing design doubled the output. 

The transmitting antenna (Figure 29) consisted of 
five dipoles coaxially mounted and backed by a plate 
reflector. This array yielded a distribution, in a plane 
containing the dipoles and normal to the reflector, 20 
degrees wide to the half-power points. In a plane 


Figure 22. Photograph of Block X AM receiver. 
































































116 


TELEVISION 


normal to the dipoles, the distribution pattern was 
120 degrees wide to the half-power points. Figure 29 
shows the complete transmitter and its antenna. 



Figure 24. Block X FM transmitter. 


The Block XVIII receiver was a superheterodyne 
with a GL-466-A lighthouse local oscillator and a 
crystal converter. The cavity of the GL-466-A had 
built-in feedback between the anode and cathode 
resonators to sustain oscillation. The resonators were 
tunable over a 150-mc range, which was arbitrarily 
divided into 10 channels of 10-mc width. 

The antenna and local oscillator were capacitively 
coupled to the converter, which consisted of a stand¬ 


ard radar crystal cartridge in a quarter-wavelength 
cavity (Figure 30). The output of the converter was 
50 me, with a 15-mc swing for the video and syn¬ 
chronizing signal. 



Figure 26. Block X FM receiver. 

The 50-mc i-f signal was amplified through nine 
pentode stages using 1852 tubes. The last two stages 
were limiters to clip noise and remove the amplitude 
modulation which produced the desired frequency 
modulation. A 6HG discriminator was 20 me wide, 
compared with the maximum frequency swing of 15 
me. This resulted in a loss of efficiency—maximum 



Figure 25. Block diagram of Block X FM receiver. 




















































SUMMARY OF RESULTS 


117 



Figure 27. Comparison of AM and FM reception: A, AM, 33^ miles; B, FM, 13^ miles; C, AM, 10^ miles; D, FM, 
5 miles. 


video output from the discriminator, 0.1 v—for 
the sake of preserving linearity of discriminator 
action. 

Video and synchronizing signals were separated 
after the second video amplifier stage. AFC by means 
of phase discrimination with the received synchroniz¬ 
ing pulses could be applied optionally to the receiver 
scanning oscillators. 

The antenna (see Figure 31) consisted of five co¬ 
axial dipoles. This array gave an omnidirectional pat¬ 
tern in a plane normal to the dipoles. In the perpen¬ 
dicular planes the sensitivity was down 3 db along the 
20-degree lines. 

Test results showed that reliable picture transmis¬ 
sion could be obtained without multipath interfer¬ 
ence if the received signal was adequate to give good 
limiter action. In stationary ground-to-ground trans¬ 
mission, limiting action was satisfactory at 6 miles 
(grazing line of sight) with a 2-watt transmitter out¬ 
put. With a 4-watt output and a 1,200-ft effective 


elevation of the transmitter, the limiter action was 
satisfactory at a 12-mile transmission distance. 

Plane-to-plane transmission yielded usable signals 
up to 15 miles with 6 watts radiated from the trans¬ 
mitter. With revised transmitter plumbing, the ra- 



Figure 28. Block diagram of Block XVIII transmitter. 


diated power was increased to between 10 and 12 
watts. Plane-to-ground and plane-to-plane tests with 
the increased transmitter power were not made. Mul¬ 
tipath difficulties were absent when the transmission 
was short enough to insure limiter action and when 



















118 


TELEVISION 


the receiver supply voltage (28 volts) was normal. 
Reduction of the receiver supply voltage to 24 volts 
produced bars on the received picture similar to 
Block III multipath patterns. Study of the i-f plate 
and screen supply was not neglected. It is perhaps 
not impossible that the reduced voltage cut the i-f 
gain sufficiently to prevent limiter action. 

A miniature Block XVIII, 21 utilizing the miniature 
Image Orthicon, was under development as World 
War II closed. The possibility of obtaining high di¬ 
rectivity from a compact antenna array suggested 
that television at Block XVIII frequencies might 
eliminate the doppler-effect difficulties with Roc. The 
project was terminated at the cessation of hostilities. 


5 5 5 Television in Missiles 

Preliminary Tests with Glide Bombs 

The glide bombs of the National Bureau of Stand¬ 
ards (see Chapter 1) were planned for remote radio 
control with television. The missile consisted of a 
monoplane airframe with a 12-ft wingspan and a 
2,000-lb GP bomb as payload. The empennage struc¬ 
ture was fixed, “guiding” being performed by deflec¬ 
tion of full-span trailing-edge wing flaps (elevons). 
Deflection of the flaps changed the camber of the 
wing, with resultant change in its lift coefficient. 
Thus the missile flew with substantially zero change 
in angle of attack with changing glide path (see tabu¬ 
lation in Section 1.3). Turns were effected by differen¬ 
tial elevon displacement. 

Tests were made of Block I in Robin (Figure 32) 
in April and July 1943 22 23 In the April tests, trans¬ 
mission was from the missile to a ground station 
where the radio-control transmitter was located. For 
such drops as had a very good initial launching, the 
received picture was adequate to permit recognition 
of the target at moderate range, about % to 1 mile. It 
was by no means good enough to permit recognition 
of the target area during early phases of the flight, 3 
to 5 miles. Therefore flights failed in which the com¬ 
bination of launching error and automatic-pilot oper¬ 
ation resulted in a heading error in excess of some 15 
degrees at a 1-mile range. Those tests whose unguided 
flight brought the missile within a mile of the target 
and at a heading which brought the target within 
the field of view of the camera scored misses of from 
100 ft to 500 ft. The ground radio-control and tele¬ 
vision-receiver station was located about 1,500 ft 


from the target. Consideration for the physical safety 
of high-ranking observers may have interfered with 
the ability of the controlling bombardier to obtain 
better scores. 

With a ground control station the conditions for 
television are somewhat better than would exist in a 
tactical situation (Figure 33). Any ground reflections 
will be directed away from the television receiver, 



Figure 29. Block XVIII transmitter and antenna. 


eliminating multipath problems. Further, the ab¬ 
sence of severe vibrations, which are inescapable in 
tactical aircraft, prevent the appearance of micro- 
phonic noise generated in the receiver. 

In the July experiments the receiver and control 
station were airborne in the bombardment airplane. 
The nose housing of the missile had been lined with 
Ozite to absorb acoustic vibrations, which were be¬ 
lieved to have impaired seriously the television per- 





SUMMARY OF RESULTS 


119 



Figure 30. Block diagram of Block XVIII receiver. 


formance in the April tests. As a further guard against 
microphonic disturbance originating in the camera- 
transmitter equipment, the camera-transmitter was 
covered with acoustic deadening material and sponge 
rubber. Microphonics were much reduced, and resi¬ 
dual disturbance was attributed to microphonics 
originating in the receiver. The total disturbance was 
sufficiently severe to cause tearing of the picture. It 
is not clear whether this interference was due wholly 
to microphonics or, at least in part, to multipath 
beating. 

At the conclusion of these experiments the Divi¬ 
sion’s work with television-guided glide bombs was 
terminated. The Navy was much more interested in 
the application of radar-homing control to this mis¬ 
sile—Pelican and Bat (see Chapter 1). The AAF had 
already under way a glide-bomb program of its own, 1 
one version of which was television-guided. The Divi¬ 
sion served cooperatively on the AAF program and 
supplied, through its contractor (RCA) under Con¬ 
tract OEMsr-441, consultation services on television 
aspects of the project. 24 The problem of guiding suc¬ 
cessfully a television-guided missile has already been 
touched on. It was particularly acute in the AAF 
glide bomb GB-4. This missile has a conventional 
empennage which provides two-axis steering, as in 


airplane practice. Consequently it flies with a vari¬ 
able angle of attack, and bore-sight errors exist at all 
but a single elevator setting. 



Figure 31. Block XVIII receiver and antenna. 









































































































120 


TELEVISION 


Television in Dirigible High-Angle Bombs 

The Division’s predecessor groups had made con¬ 
tracts with Hazeltine Service Corporation, 25 > 26 Farns¬ 
worth Television and Radio Corporation, 27 and 
RCA 7 ' 14 * 28 ’ 29 to develop compact television pickup and 
transmitting equipment of such compass as to be 
contained within a standard aircraft bomb. None of 
the contractors succeeded in producing a reliable 
assembly which would be contained within a 1,000-lb 
GP bomb and leave space for an appreciable charge 
of explosive. The equipment would fit in the compass 
of the standard 2,000-lb bomb with space for ap¬ 
proximately the bursting charge of a 1,000-lb GP 
bomb. 

The Hazeltine project compromised resolution and 
flicker as a concession to compactness. To obtain 
minimum bulk the scanning frequency was reduced 
to 210 lines per frame and 10 frames per second. 
Transmission on a 220-mc carrier was by frequency 
modulation for the video signal and by amplitude 
modulation for the synchronizing and blanking 
pulses. The radiated output was approximately 1 
watt. 

To compensate for the large variation in angle of 
attack which a high-angle dirigible bomb undergoes, 
vanes mounted in the wind stream were coupled to 


the pickup assembly so as to keep the camera objec¬ 
tive continuously aligned with the flight path. 

Experiments indicated that the resolution and 
power output were inadequate and that flicker was 
at an objectionable level. 

The Farnsworth project employed the Image Dis¬ 
sector. Its comparative freedom from microphonics 
(see Figure 8) and circuit simplicity—thirteen tubes 
and four controls in the Dissector camera as compared 
with seventeen tubes and fourteen controls for the 
Block III Iconoscope—made it attractive for missile 
application, although its low sensitivity was a draw¬ 
back. Transmission was at 112 me, amplitude modu¬ 
lated with a single 4-mc sideband on the basis of 225 
lines and 40 frames sequentially scanned. 

The transmitter oscillator consisted of two 6C4 
triodes in a modified Hartley circuit. The oscillator, 
plate modulated by the video and synchronizing- 
signals, was transformer coupled to the power ampli¬ 
fier, which consisted of an 832 twin pentode in push- 
pull. This circuit was able to radiate approximately 
10 watts. Plane-to-plane tests produced a received 
picture with bars characteristic of spurious frequency 
modulation. The picture was, however, deemed ade¬ 
quate to justify drop tests. Poor weather frustrated 
the program, and dropping tests were never carried 
out. Figure 34 shows nose and tail views of a dirigible 



RADIO 

RECEIVER. 


TELEVISION 

CAMERA 

TRANSMITTER 

UNIT 


Figure 32. Robin, nose fairing removed. 





SUMMARY OF RESULTS 


121 



bomb C equipped with Farnsworth television cam- 
era-transmitter equipment. 

The Vericon equipment (Figure 35) was manufac¬ 
tured by Remington Rand. The camera tube was 
similar to the Orthicon of RCA (see Figure 14). Its 
development was undertaken by the contractor be¬ 
cause the difficulties in the control of secondary elec¬ 
trons in pickup tubes of the Iconoscope type seemed 
insurmountable to them. 

The 105-mc carrier was amplitude modulated with 
video and synchronizing signals. Scanning frequency 
was based on 350 lines and 30 frames per second. The 
master oscillator was crystal-controlled at 26.25 me 
and was followed by two frequency doublers in a 
6N7. The video and synchronizing signals grid modu¬ 
lated the power amplifier, two 829 tetrodes in push- 
pull. The power output was approximately 10 
watts. 

The resolution obtained by the contractor was 
rather better than that obtained by Columbia in the 
comparative study of pickup equipments (Figure 36). 
The reason for this has never been made wholly clear. 
The automatic beam-focusing control included in the 
camera circuit may have been a contributing factor. 
As in the case of the Farnsworth equipment, drop 
tests were planned for this equipment but were can¬ 
celled on account of poor weather. 

Drop Tests. In addition to the foregoing, which 
were never conclusively tested, the Division had a 
project with RCA to adapt the Block 1,100-mc, equip¬ 
ment to the guidance of the high-angle dirigible 
bomb (Figure 37). Preliminary tests were held at 
Eglin Field in the winter of 1942-43. They showed 
sufficient promise to justify continuing development 
work. Further tests with revised television equip¬ 
ment were made at Tonopah Air Base in April and 
May 1944. While the results as regards television 
were probably satisfactory, the tests emphasized how 
different the problem of steering such a missile was, 
even with wholly satisfactory television performance. 
In view of the Division’s program with Roc (see 
Chapter 4) and Mimo, work on a television-guided 
high-angle bomb was shelved. 

The Eglin Field experiments comprised drops with 
three missiles. One was a dummy, provided with a 
parachute which extended the time of fall from ap¬ 
proximately 30 seconds to 6 minutes; this permitted 
more complete observations of missile-to-plane trans¬ 
mission than could be made with an unimpeded drop. 
This dummy missile was not equipped either with 
radio control or gyro stabilization. The other two 


missiles were laboratory prototypes to explore the 
possibility of accurate steering. 

Preliminary tests to determine radiation strengths 
and approximate patterns were made by standing the 
missile on end on the ground and transmitting a video 
signal to an airplane 15,000 ft above it. These tests 


Figure 33. Photographs of received picture at ground 
control station: A, 1 min before impact; B, 30 sec be¬ 
fore impact; C, 10 sec before impact. Arrows indicate 
target. 






122 


TELEVISION 


disclosed that the radiation core from the missile was 
approximately 37 degrees wide to an undefined mini¬ 
mum signal-strength level. They also disclosed a 
serious pulse-modulated interfering signal at almost 


exactly the carrier frequency of the Block I trans¬ 
mitter. This proved to be a nearby radar search sta¬ 
tion operating at 105 me. Further work was carried 
forward by an informal agreement to shut the radar 



PAY LOAD (OR BALLAST) 
MOUNTED IN CENTER 


NOSE SECTION 


WITH PLASTIC 
WINDOW, MOUNTS 
HERE (NOT SHOWN) 


HAND HOLE (FOR SERVICING) 


ANTENNA 

STRUCTURE 


POWER SUPPLY a BATTERIES 
MOUNTED ON BACK RING 


CONVERSION-XMITTER 
SHOCK-MOUNTED TO FRONT RING 



CONTROL CIRCUITS 
MOUNTED IN 
TAIL 


REFLECTORS 


ANTENNA MATCHING 
CIRCUITS 


ANTENNA CABLE 


RADIATORS 


AILERONS 


Figure 34. Nose (fairing removed) and tail views of dirigible bomb equipped with Farnsworth television camera-trans¬ 
mitter equipment. 







SUMMARY OF RESULTS 


123 


station down when it was desired to put the television 
on the air. 

The dummy was dropped with parachute recovery. 
Picture reception was considered satisfactory, al¬ 
though there was some loss of horizontal synchro¬ 
nism. This the television operator at the receiver was 
able to correct promptly. The equipment having been 
recovered undamaged, the dummy was then dropped 
unimpeded. The trouble with horizontal synchroniza¬ 
tion continued but was deemed inconsequential. 




Figure 36. A, standard resolution chart; B, picture re- 
Figure 35. Vericon camera installed in dirigible high- ceived from Vericon pickup, 

angle bomb. 


The two prototype missiles were then dropped. The 
enthusiasm of the investigators was undampened, in 
spite of a report which stated, “There was a very ob¬ 
jectionable white band about an inch wide [out of a 
viewing screen about 4 inches wide) and two inches 
down from the top which extended completely across 
the picture raster.” This effect, which had been ob¬ 
served before, was most probably caused by inter¬ 
ference with the ground-reflected signal, inasmuch as 
it was not observed at the ground monitor station. 
Other difficulties included: (1) horizontal bars due to 
crosstalk with the radio-control carrier, or possibly 
to microphonics; (2) vertical bars probably due to 
multipath; (3) loss of horizontal synchronism due to 
fluctuation in the battery voltage, as the coverage of 
the plates by electrolyte varied with changes in the 
missile attitude. 



Figure 37. Television-guided high-angle bomb. 

















124 


TELEVISION 


The Tonopah experiments were made with Block I 
equipment in which the improvements cited in Sec¬ 
tion 5.5.1 had been incorporated. Clamp circuits were 
included in the camera-transmitter units. All receiv¬ 
ers had a tunable r-f head and low-impedance AGC. 
Receivers planned for use in controlling the missile 



Figure 38. Folded-dipole antenna on television- 
guided high-angle bomb. 



Figure 39. Receiver in B-25 airplane showing arrange¬ 
ment of compensating crossed wires on television screen. 


had phase-actuated AFC. An antenna of improved 
gain to increase the radiated power was installed on 
the bomb (Figure 38). To compensate for varying 
angle of attack with application of control, stiff wires 
were mounted vertically and horizontally in front of 
the viewing screen (Figure 39). With no control signal 
applied, these wires crossed in front of the center of 
the screen; upon application of control signal, they 
moved across the screen horizontally and vertically 


at a speed proportional to the rate of rudder or eleva¬ 
tor displacement. To a first-order approximation (see 
Chapter 2), this action compensated for changes in 
angle of attack during steering, and the point of in¬ 
tersection of the crossed wires corresponded to the 
point on the transmitted image of the terrain toward 
which the missile was heading. To avoid crosstalk be¬ 
tween the control and television signals, the trans- 



Figure 40. Radio-control transmitting antenna. 



Figure 41. Television receiving antennas. 


mitting antenna (Figure 40) for control signals was 
mounted in front of the bombardier’s greenhouse. 
The television receiving antennas (Figure 41), two 
parallel dipoles, were mounted under the belly of the 
ship. 

In addition to the receivers in the plane, a group 
monitoring station was installed in a blacked-out 23 ^- 
ton truck near the bombing range. Reliable radio 













SUMMARY OF RESULTS 


125 


communication was maintained between the ground 
monitoring station, the airplane, and the target area. 
The target consisted of scraped areas on the desert 
floor—a 100-ft diameter bull’s-eye with scoring rings 
of 200-ft and 500-ft diameter. To improve the low 
contrast between the scraped target and the general 
terrain, slaked lime was spread on the bull’s-eye and 
scoring rings. This did not, however, remain in place 
long on account of the high winds prevailing at that 
season. 



Preliminary to the drop tests, passage tests were 
made in which all signals which might be on the air 
during a drop were simultaneously operated. There 
was no crosstalk, and the received television signal 
was reasonably free from microphonics. No picture 
was received since the bomb ,was enclosed within the 
bomb bay of the B-25, and the glass window of the 
bomb was masked. 

In the first drop the operator never saw the target, 
and the miss was 2,000 ft in range and 2,800 ft in 



Figure 42. Received pictures in airplane: A, B, and C, time of release and altitude not known; D, 21.8 sec from release, 
altitude 3,860 ft; E, 28.1 sec from release, altitude 1,035 ft; F, 28.8 sec from release, altitude 250 ft. 






126 


TELEVISION 


azimuth. In the remaining four drops the errors were: 


Drop No. 2 
Drop No. 3 
Drop No. 4 
Drop No. 5 


Range error Azimuth error 
300 ft 250 ft 

No record 

450 ft 525 ft 

15 ft 250 ft 


It was the consensus that television can produce an 
adequate picture of the terrain to improve bombard¬ 
ment materially during phase 3 (see Section 5.1) of 


an attack. The* coupled crossed wires, however, are 
inadequate to compensate for variations in angle of 
attack. Bars due to interference between the signal 
from the missile and that from its image (see Figure 
5) is a problem as yet unsolved. These bars typically 
appear about 15 seconds after release, when the bomb 
velocity has reached some 450 ft per second. They 
are shown in Figure 42, which is a group of single 
16-mm frames from the motion-picture record of the 



Figure 43. Received picture at ground monitoring station: A, 26 frames, 1.03 sec from release; B, 48 frames, 3 sec 
from release; C, 198 frames, 12.3 sec from release; D, 241 frames, 15.06 sec from release; E, 450 frames, 28.1 sec from 
release, altitude 1,035 ft; F, 461 frames, 28.8 sec from release, altitude 250 ft. 













SUMMARY OF RESULTS 


127 



receiving screen in the aircraft. Their absence from 
the pictures received on the ground (Figure 43) is 
striking in comparison. 

Even if the interference were not sufficiently severe 
to threaten to obscure the target and with adequate 
compensation for variation in angle of attack, the 
problem of steering is still paramount. If the missile 
is so maneuvered as to place the target early in the 


Figure 44. Mimo camera (right) compared with 
standard Block III camera (left). 


Figure 46. Mimo transmitter, side view, covers re¬ 
moved. 

4 of this report, recommended itself to television 
guiding during phase 3 of an attack for several rea¬ 
sons. It was designed to fly with the axis of its fuse¬ 
lage continuously tangent to the flight path—zero 
angle of attack—so that the center of a television 
image received from a camera-transmitter bore- 


Figure 45. Mimo camera, cover removed. 


center of the receiver screen it will be impossible to 
hold it there, as the bomb will have been dived too 
early and adequate lift is not available with the high- 
angle dirigible bomb to pull out at the end. A lead 
computer which takes competent account of the dy¬ 
namic trajectory of the missile is indicated. The de¬ 
velopment of such a computer for the high-angle 
dirigible bomb seems unjustified. 

It had been planned to test the Image Dissector and 
the Vericon equipments during theTonopah program. 
The onset of a long period of bad flying weather and 


the rather definite contra-indications in the results of 
the Block I experiments recommended a decision 
against further work with the television high-angle 
bomb. 

Roc and Mimo Project 21 

The missile Roc, developed for the Division by 
Douglas Aircraft Company and discussed in Chapter 















128 


TELEVISION 


sighted with the axis of the missile would continu¬ 
ously indicate the point of impact of the missile with 
continued rectilinear flight. The normal trajectory 
was relatively steep—intermediate between the dir¬ 
igible high-angle bomb and the glide bomb—so that 
the received picture of the terrain showed little fore¬ 
shortening. Finally, its maneuverability, which was 
7,500 ft minimum turning radius as compared with 
20,000 ft for the production Razon, seemed to justify 
the use of television as a means of guiding, costly in 
complication and manpower though it was. Accord¬ 
ingly the Division, strongly urged by the Air Tech¬ 
nical Service Command, undertook the development 
of a compact camera and transmitter equipment 
specifically for this missile. Miniature in compass and 
utilizing an Image Orthicon-type of pickup tube, it 
was dubbed Mimo. Transmission was set in the 300- 
mc band, although that frequency had been pre¬ 
empted by other communication services. The hope 
of getting a compact transmitter and carrier at 800 
or 1,800 me in time to be of military significance 
seemed remote. 

Camera. The camera utilized a new tube similar in 
design and performance to the Image Orthicon (see 
Figure 16). It was, however, considerably smaller, 
being but 9 in. long overall and l}/£ in. in diameter 
(Figure 44). All leads were brought out to a 17-prong 
base. This new design was brought out rapidly and 
smoothly. Except for loss of sensitivity beyond 0.71 ^ 
on account of contamination of the photocathode by 
gas driven off from the electron cathode, its perform¬ 
ance was identical with the larger version. 

The entire camera unit (Figure 45) contained a 
video amplifier, deflection, and synchronizing cir¬ 
cuits. The video amplifier contained four stages of 
6AK5 miniature pentodes. The third stage had a 
high-frequency peaking circuit in the grid input to 
provide substantially flat response out to 4 me, in 
spite of the high-frequency attenuation of the Image 
Orthicon. Beyond the fourth stage, a 6AK5 clipper 
provided a 0.3-v blanking pedestal for the synchro¬ 
nizing signal. 

Both deflection circuits consisted of 3A5 twin tri- 
odes, operating as blocking oscillators and discharge 
tubes. The vertical deflection 40-c oscillator was fol¬ 
lowed by a 3A5 amplifier output stage with both sec¬ 
tions connected in parallel; the horizontal deflection 
oscillator was followed by a 25L6 output stage. In 
addition to energizing the deflection coils for the pick¬ 
up tube, each deflection output tube drove a blanking 
stage. The output of the horizontal blanking stage 


supplied a voltage to the clipper to create the blank¬ 
ing pedestal; the output of the corresponding vertical 
stage was applied to the target of the pickup tube to 
provide vertical blanking. The output voltages of the 
deflection circuits were mixed in a twin triode to pro¬ 
vide the synchronizing signal. 

Transmitter. The entire tube complement of the 
r-f section of the transmitter (Figure 46) consisted of 
three 2C43 lighthouse tubes. The master oscillator, 
which operated at carrier frequency, was tuned with 
a resonant line and had Colpitts feedback. An adjust¬ 
able feedback capacitor determined the plate current 
at the proper level. The output of the oscillator was 
impressed, through resonant lines, on the grids of 
two 2C43’s in push-pull, which comprised the output 
power amplifier. 

The video and synchronizing signals were mixed at 
the first-stage grid of a three-stage video amplifier. 
The output stage, two 6V6’s in parallel, grid modu¬ 
lated the power amplifier. When connected to a 
matching antenna, the transmitter radiated 7 to 10 
watts. 

Antennas. Four antennas were required for the Roc 
system. A transmitting antenna was necessary for the 
television equipment on the missile; this was known 
as Mimo-Roc. A receiving antenna for the channel 
was required at the plane; this was known as Mimo- 
Plane. A receiving antenna on the missile was re¬ 
quired for the 84-mc control—Roc; a complementary 
transmitting antenna on the plane was known as 
Control-Plane. The Mimo-Roc antenna had to be so 
designed as to give a radiation pattern which would 
contain the plane at a bearing of high signal strength 
for any expected maneuver of the plane or missile. 
It also had to cut off forward radiation sharply so as 
to prevent, as far as possible, any signal from reach¬ 
ing the ground to produce multipath, as well as to 
conserve power. The elimination of downward radia¬ 
tion is tactically important to preserve security as 
well as being technically important for the reasons 
just mentioned. The Mimo-Plane transmitter ideally 
should cut off sharply beyond the limits of the cone 
originating in the plane which will contain the missile 
for all practicable maneuvers. The elimination of 
downward sensitivity of the Control-Roc antenna is 
an aid against jamming. The requirements of the 
Control-Plane antenna are similar to those of the 
Mimo-Plane. 

The Mimo-Roc antenna (Figure 47) consisted of a 
tuned dipole mounted at the rear of the missile, with 
the axis of the dipole arms parallel to the pitch axis 





SUMMARY OF RESULTS 


129 


of the bomb. The Roc was gyrostabilized in roll by 
means of ailerons. Therefore its pitch axis is continu¬ 
ously parallel to the ground plane. The radiation pat¬ 
tern from the dipole can then be defined in terms of 
two orthogonal planes: the azimuth plane is one 
which contains the axis of the dipole arms and the 



Figure 47. Antennas on Roc: A, Control-Roc receiv¬ 
ing antenna; B, Mimo-Roc dipole. 


RELATIVE FIELD INTENSITY 



BEARING IN DEGREES 

Figure 48. Radiation pattern of Mimo-Roc dipole. 


ness. The problem, then, was to ground the dive 
brake to the structure at 300 me but to insulate it 
therefrom at 84 me. This was accomplished by build¬ 
ing the supporting struts of composite Dural-Bake- 
lite. This construction provided very low capacitive 
reactance at Mimo frequencies but kept it high at 
control frequencies. 



Figure 49. Slot antenna for Mimo-Roc with folded 
waveguide. 

RELATIVE FIELD INTENSITY 



BEARING IN DEGREES 

Figure 50. Distribution pattern of Control-Roc 
antenna. 


axis of its stem; the elevation plane is perpendicular 
to it and also through the axis of the dipole stem. 

It was desired to use the perforated dive brake as 
a reflector for the dipole. Its size and location were 
determined by aerodynamic arguments principally to 
give zero angle of attack at trim. It was also desirable 
to use the same element as the Control-Roc antenna 
in view of the need to maintain aerodjmamic cleanli¬ 


With this arrangement, less than 10 per cent of the 
maximum signal strength was radiated in a forward 
(toward the ground) direction. (See Figure 48.) The 
width of the useful beam in the azimuth plane was 90 
degrees to the half-power points; in the elevation 
plane the width to the half-power points was 125 
degrees. 













130 


TELEVISION 


An alternative antenna design was completed which 
consisted of a slot driven by a folded waveguide 
(Figure 49). The distribution from this antenna was 
not materially better than from the tuned dipole, and 
it had the disadvantage of being extremely critical to 
carrier frequency. 



Figure 51. Mimo-Plane antenna array mounted on 
test panel. 


RELATIVE FIELD INTENSITY 



BEARING IN DEGREES 

Figure 52. Horizontal distribution pattern of Mimo- 
Plane antenna. 


The Control-Roc antenna consisted of the dive 
brake. Its sensitivity pattern (Figure 50) was not 
ideal but it was deemed thoroughly satisfactory for 
tests and probably adequate for initial combat use. 

The Mimo-Plane antenna consisted of two quar¬ 
ter-wavelength dipoles (Figure 51) mounted under 
the skin of the ship. The dipoles were so placed as to 


straddle the plane of the ship’s yaw and roll axes. 
Their centers were spaced approximately one-half 
wavelength, and the dipoles were oriented so that 
their extended axes would intersect at 90 degrees and 
on the ship’s yaw-roll plane. Figure 52 shows the 
omnidirectional character of this array in the hori¬ 
zontal plane; the distribution in the other planes is 
shown in Figures 53 and 54. 


RELATIVE FIELD INTENSITY 



BEARING IN DEGREES 

Figure 53. Vertical-transverse distribution pattern of 
Mimo-Plane antenna. 

RELATIVE FIELO INTENSITY 



Figure 54. Vertical-longitudinal distribution pattern 
of Mimo-Plane antenna. 


The Control-Plane antenna was the standard di¬ 
pole developed for Razon. 

Results of Tests. Ten Roc missiles were dropped 
with Mimo. The results of the first six drops, which 
are tabulated below, indicated more promise for this 
system than any other system of television-guided 
missile that the Division has encountered. 

Drop T-l Bad picture due to defective pickup tube. Prob¬ 
ably spots on photocathode or target. 

Drop T-2 Usable picture all the way, somewhat marred by 
doppler multipath after about 15 seconds of flight.. Miss: 
about 75 ft. 














SUMMARY OF RESULTS 


131 


Drop T-3 Usable picture until the last 10 seconds of flight . 
Very severe multipath developed so an overshoot was cor¬ 
rected near the end of the flight. Miss: about 300 ft. 

Drop T-4 Spurious oscillation in transmitter output am¬ 
plifier; also serious multipath. No guiding was practicable. 

Drop T-5 Picture faded because the receiver had been tuned 
to the carrier image. The receiver was retuned; the balance of 
the flight—only a few seconds—was marred by multipath. 
Miss: about 220 ft. 

Drop T-6 The target failed to appear in the field of the 
missile on account of an excessive crab angle in the ship at the 
instant of release. As the picture of the terrain appeared, it 
was marred by very serious multipath interference. 

Although the target itself (Figure 55), had been 


scraped clear of desert vegetation and covered with 
several tons of salt crystals, contrast was, in general, 
inadequate (Figure 56). The desert floor beyond the 
target area was extremely deficient in contrast, so 
that when the target was once lost it was virtually 
impossible to regain it. 

As a result of the foregoingi experiments serious ef¬ 
forts were made to improve the contrast of the tele¬ 
vised picture and to eliminate the doppler multipath 
effect. A search for a better Mimo-Roc antenna at 
this frequency was fruitless. Faster AGC was applied 
to the receivers in the hope of mitigating the effects 



Figure 55. Wendover target range for Roc tests. 



132 


TELEVISION 


of doppler multipath interference. A selection of the for the remaining missiles. To improve contrast, the 
best Mimo tubes from the stock available was made / 2.0 lenses were replaced with lenses of / 4.5 aper- 



Figure 56. Photograph of television screen during drop T-2. 












SUMMARY OF RESULTS 


133 


ture; all optical surfaces were treated with fog- 
repellent coatings. 

With the foregoing revisions four more drop tests 
were made. 

Drop T-7 Horizontal synchronism was lost for about 134 
seconds. Doppler multipath bars appeared for the last half of 
the flight but their intensity was somewhat relieved by the 
fast AGC. Miss: about 265 ft. 

Drop T-8 Substantially the same as T-4. Miss: about 
118 ft. 

Drop T-9 Some doppler multipath during the last half of 


the flight. The missile was controlled from the ground monitor 
station where the received picture was not marred by multi- 
path interference. Miss: about 140 ft. 

Drop T-10 Roll stabilization lost. 

Guiding by means of a lead-computing aid (see 
Chapter 4) was not attempted as the termination of 
the Division’s activities intervened. The work is being 
carried forward by the AAF. As this report is pre¬ 
pared, word has been received that at least one drop 
made by the Army has landed well within a 50-ft 
bull’s-eye circle from 15,000 ft. 



Chapter 6 

RADIO-CONTROL SYSTEMS 


INTRODUCTION 

T he radio problem confronting the Division at its 
formation was threefold: (1) it was necessary to 
obtain remote-control radio links satisfactory for the 
control of the missiles during their developmental 
stages; (2) radio systems adequate for the initial 
combat phase had to be made available as rapidly as 
missiles requiring their use were ready for operation; 
(3) finally, there was a responsibility to maintain a 
continuing program, seeking to develop systems of 
steadily increasing security in order that new and less 
jammable systems would arrive in the theater before 
the enemy had developed methods of jamming those 
already in operation. 

It would seem that such a problem could hardly be 
acute. No nation was ever more “remote-control con¬ 
scious” than the United States; no other people had 
developed radio communication to anything like an 
equal degree. It would appear to be necessary simply 
to procure suitable radio links of already established 
designs from the sources of proved reliability. Such 
was the approach by the Division. It proved a gross 
error. 

No mistaken judgment plagued the activities of the 
Division more than its failure to realize the difficulty 
of the radio-control problem and to meet it head on. 
The problems were both organizational and techno¬ 
logical. All electronic devices for Army use were the 
responsibility of the Signal Corps; guided bombs were 
the responsibility of the Army Air Forces so long as 
the actual metal container for the explosive was not 
altered, when the Ordnance Department became an 
interested party. Without the assistance of the Air 
Communications Officer to resolve questions of 
divided responsibility, the whole radio-controlled 
bomb program within the Division might well have 
failed. 

On the technological side many new problems of 
unattended operation under conditions of extremes 
of temperature, humidity, altitude, vibration, and 
acoustic perturbation had to be solved. Further, in 
the transmitting equipment, which was by nature 
airborne, weight was at a premium; the space limita¬ 
tion on the receivers and the aerodynamic limitations 
on their antennas were extreme. 


62 RADIO SYSTEMS FOR TEST 
6,2,1 Radio Systems for Glide Bombs 

In connection with early tests of Robin (see Chap¬ 
ter 1) the National Bureau of Standards secured 
through Contract NDCrc-141 a small quantity of 
standard police radio receivers and a transmitter 
from RCA. These were modified by RCA so that, 
when a control impulse was transmitted, the gyro 
bias coils received a fixed value of current to produce 
a fixed rate of turn or a fixed rate of change of glide 
path. (See Section 1.6.) Space was adequate within 
the Robin for a standard receiver, and the wings 
formed a convenient surface to which a thin ribbon 
antenna could be cemented without producing any 
aerodynamic disturbance. 

6 2 2 Radio Systems for High-Angle Bombs 

For Azon and Razon, however, the situation was 
entirely different. No space could be made available 
in the bomb which would accommodate a standard 
receiver. Development was required, but the usual 
sources of electronic developmental skill were bur¬ 
dened with other war work of high priority and were 
hardly to be persuaded to undertake development 
programs of the speculative character which guided 
missiles had in their beginnings. 

At the outset the Gulf Research and Development 
Company considered essential a system in which 
the rudder and/or elevator displacement was pro¬ 
portional to the stick displacement. This later 
proved erroneous (see Section 2.9 and Chapter 10), 
but the principle was adhered to throughout the 
early work. In their search for a commercially avail¬ 
able radio, Gulf found two systems. The first had 
been developed by Bendix Aviation for the AAF 
target and glide-bomb program; the second was 
developed by the American Junior Aircraft Com¬ 
pany 1 for the control of model airplanes. The Ben¬ 
dix system provided on-off control rather than 
proportionality. Furthermore, when control was ap¬ 
plied in two components, yaw and pitch simul¬ 
taneously, there was crosstalk between the audio 
channels. This system was, therefore discarded by 
Gulf. 


134 


RADIO SYSTEMS FOR TEST 


135 


The principle of the American Junior Aircraft 
Company was adopted. Its embodiment was, how¬ 
ever, deemed too complicated for satisfactory ex¬ 
perimental work on bombs. The transmitter radiated 
two unmodulated CW signals at 84 me and 86 me; 
one frequency controlled yaw, the other pitch. The 
CW signals were cyclically keyed so that pulses of 
each were radiated. The position of the control stick 
determined the ratio of pulse length to the keying 
cycle. Equal time-off and time-on indicated a neutral 
rudder or elevator position. 



Figure 1. Block diagram of American Junior A/C 
system of radio control. 


The receiver was a crystal-controlled superhetero¬ 
dyne (Figure 1). An antenna stage was spot-tuned 
to the two frequencies. Two i-f strips, detectors, and 
d-c amplifiers accommodated the azimuth and range 
channels. The output of each d-c amplifier was a 
series of pulses of varying length with constant total 
period. The last stage of the d-c amplifier was driven 
from cutoff to saturation so that the pulses were of 
constant amplitude. These pulses were fed to spring- 
restrained miniature torque motors. The inertia of 


space the r-f end of the receiver was changed to a 
superregenerator with a separate quench oscillator 
(Figure 2). The Q of the detector-oscillator circuit 
was kept low enough so that triggering by the quench 
oscillator was required. For test purposes the broad 
acceptance of the superregenerative circuit was at¬ 
tractive, as there was but little danger of failure of a 
test due to transmitter drift or errors of receiver 
tuning. 

The control end of the system was modified by re¬ 
placing the torque motor by an electronic integrator, 
followed by a double-throw relay. With the relay 
energized, the rudder servomotors operated in one 
direction; with the relay de-energized, the motor re¬ 
versed. A potentiometer on the servomotor provided 
a follow-up voltage which biased the final stage to 
cutoff when the rudder reached a displacement corre¬ 
sponding to the control-stick position. Enough 40-c 
ripple was allowed to appear in the relay voltage to 

BIAS FEEDBACK < 


Figure 3. Block diagram of Gulf Razon receiver. 


QUENCH 

OSCILLATOR 


ANTENNA 


DETECTOR 

STAGE 


OSCILLATOR 




Figure 2. Block diagram of Gulf Azon receiver. 


the armatures with the compliance of the restraining 
springs acted as a mechanical integrator, averaging 
the pulses so that the displacement of the motor was 
proportional to the fraction of the cycle occupied by 
the CW pulse. A commutator with sectors on the 
torque motor governed the servomotors driving the 
control surfaces. 

The American Junior Aircraft Company system 
was modified by Gulf for the Azon program. The 
transmitter system was kept intact, with a keying 
frequency of 40 c. For economy of tubes, power, and 


cause relay chatter and thus prevent hysteresis in the 
cores from biasing the relay response. 

For Razon control Gulf developed and proposed to 
use a further modification of this system (Figure 3). 
The speed of keying under this system was to be 
varied as well as proportion of time-on to cycle 
length. 2 Displacement of the control stick in the azi¬ 
muth sense varied the pulse length as before; dis¬ 
placement in the range sense varied the keying fre¬ 
quency. In the receiver a discriminator stage with a 
d-c amplifier followed the a-f amplifier, in parallel 
with the integrator and its d-c amplifier. This stage 
produced a d-c voltage proportional to the keying 
frequency. A potentiometer on the elevator motor 
provided a follow-up voltage to bias the last stage to 
cutoff as before. 

This circuit was built and fully tested in the labor¬ 
atory at Gulf. It is given here complete (Figure 4) as 
it has not been elsewhere reported. 














































136 


RADIO-CONTROL SYSTEMS 


POT COUPLED 
TO RUDDER 



Figure 4. Wiring diagram of Gulf Razon receiver. 


63 RADIO SYSTEMS 

FOR INITIAL COMBAT USE 

Azon Receivers 

One of the features which made the Gulf Azon 
receiver attractive for experimental work spoke 
strongly against its use in combat. Its broad accept¬ 
ance made it most undesirable since it precluded the 
use of closely spaced channels for the simultaneous 
control of Azons from separate aircraft. It was be¬ 
lieved to be subject to microphonic disturbance; in 
the tests at Muroc on September 10,1943, the receivers 
had been packed in sand in the body of the bomb to 
avoid this difficulty. The superregenerative receiver, 
while most economical, is notoriously unstable under 
changes in temperature. Furthermore, the whole 
system of proportional control which had given rise 
to the Gulf radio development program had been 
found to be unnecessary. Even when it was available 
the controlling bombardier had successfully guided 
the bomb with rudders full-right, full-left, or neutral. 3 
Had proportional control been necessary, a different 
system would have been required as loss of the con¬ 


trol signal with this system results in full-left rudder, 
which would have produced gross misses. 

Under a general consulting-engineering contract 
(OEMsr-240), MIT made a summary study of the 
problem and recommended the use of a simple super¬ 
heterodyne receiver with a crystal-controlled local 
oscillator and a single tuned-antenna stage. The 
transmitter recommended was the existing standard 
RC-186, with provision for selectively modulating 
the carrier with six audio tones. For Azon one tone 
was selected for full-right rudder, a second for full- 
left rudder. Absence of an audio signal was to result 
in neutral-rudder position. This recommendation 
was submitted to the Army by the Division at a 
meeting called by the Air Communications Officer 
under the auspices of the Joint Chiefs of Staff. The 
recommendation was accepted without dissent. In 
spite of such expressed unanimity, the Signal Corps 
liaison officer on the project, in whom responsibility 
for the selection rested, insisted on the use of a super- 
regenerative receiver. In the face of such inflexibility 
the Division was left without option. 

A contract (OEMsr-1081) had been negotiated b}^ 
















































































































RADIO SYSTEMS FOR INITIAL COMBAT USE 


137 


the Division with Union Switch and Signal Companj^ 
to engineer Azon for production. This contractor 
now made a subcontract with General Instrument 
Company of Elizabeth, New Jersey, to develop a 
superregenerative receiver 4 which would operate with 
the RC-186 transmitter in the manner planned for 
the superheterodyne. This became accepted as the 
A/N-CRW-2A (Figure 5). In accepting it, the Signal 
Corps officer found it necessary to waive many ap¬ 
plicable portions of Specification No. ARL-102-A. 
This receiver was produced in some quantity under 
subsequent Army contracts. A similar receiver, on the 
whole rather less satisfactory (the A/N-CRW-2), 
was produced in even greater volume by a different 
contractor. 

6 3 2 Razon Receivers 

The unsatisfactory experience with the Azon re¬ 
ceiver taught the Division that a radio system satis¬ 
factory for initial combat must be available before 
the Razon would be ready for production. Accord¬ 
ingly, MIT extended their studies under the guidance 
of the radio specialist in the Division. 5 This more 
thorough study confirmed their summary recom¬ 
mendations in favor of a crystal-controlled super¬ 
heterodyne receiver. They further recommended an 
i-f amplifier operating either below 300 kc or one op¬ 
erating between 10 and 20 me. The lower frequency 
held the possibility of resistance coupling, with con¬ 
sequent simplicity and stability; the higher frequency 
had promise of successful application to supersonic 
modulation control frequencies or to pulse techniques, 


each of which offered hope of increased security. 

Accordingly, the Division undertook both pro¬ 
grams. A contract was made (OEMsr-1195) with 
Harvey Radio Laboratory of Cambridge, Massachu¬ 
setts, for the construction, under the close guidance of 
the MIT consulting engineers, of receivers utilizing a 
75-kc i-f amplifier. Under 1 Contract OEMsr-1314, 
Philco 6 constructed receivers using a 15-mc i-f strip. 
Each group added special features although, in the 
main, the receivers were similar. The MIT-Harvey 
receiver was provided with an AGC actuated from 
the output of the detector. This type of AGC has the 
result of widening the acceptance of the receiver. 
However, if a fifth modulating tone, already avail¬ 
able in the standard RC-186 control transmitter, is 
applied whenever the stick is in neutral, it increases 
by nearly tenfold the strength of jamming signal 
required to block the control. The Philco receiver had 
a plug-in fixed-tuned r-f and local oscillator section, a 
real saving in storage and issue. Figures 6, 7, 8, and 9 
show circuits and photographs of the MIT-Harvey 
and the Philco receivers respectively. 

Figdre 10 shows the comparative selectivity of the 
A/N-CRW-2A, the Philco, and the MIT-Harvey re¬ 
ceivers. The last is shown with the audio-actuated 
AGC and with the more conventional form. For a 
sensitivity 2,000 times down (i.e., microvolts at an¬ 
tenna to operate relays is 2,000 times value at 
resonance) the band widths are: 

MIT-Harvey 
Conven¬ 
tional Audio 
A/N-CRW-2 A Philco AGC AGC 
Bandwidth More than 5 me 935 kc 450 kc 2.5 me 


DETECTOR AUDIO 

R-F STAGE OSCILLATOR AMP 



Figure 5. Wiring diagram of A/N-CRW-2A. 































































ANTENNA CONNECTOR 


138 


RADIO-CONTROL SYSTEMS 







• • 


Figure 6. MIT-Harvey receiver circuits. 



















































































































































































































































































































































RADIO SYSTEMS FOR INITIAL COMBAT USE 


139 


The sensitivities as the battery voltage declines are 
shown in Figure 11. 

The increase in bandwidth of the MIT-Harvey 
receiver, due to audio AGC, is shown in the preceding- 
paragraph. The antijam tests showed that with a 
200-microvolt signal, which is rather below the ex¬ 
pected minimum signal strength, from the control¬ 
ling plane, the enemy would have to transmit 
to the Razon an 1,800-microvolt CW signal within 
200 kc of resonance in order to block the control. 
This antijamming ratio of approximately 9 re¬ 
mains about constant with change of control-signal 
strength. 

Samples of each of these receivers were sent to the 
Aircraft Radio Laboratory for test. This organization 


made a contract with Delco Radio for the develop¬ 
ment of another Razon receiver. It is substantially 
the same as the Philco circuit. Crystal control of the 
local oscillator is retained. The fixed-tuned r-f head 
and the audio AGC are omitted. It should also be 
noted that the final stagey of this receiver, the 
A/N-CRW-7, are not biased to cutoff, which makes 
it extremely sensitive to noise generated by other 
equipment in the bomb-control assembly. 

Some fifty of the MIT-Harvey sets were built and 
used in Razon and Roc tests while the decision as to 
the A/N-CRW-7 was pending. After the manufac¬ 
turing eccentricities present only in the first seven 
units were eliminated, the sets were wholly satis¬ 
factory. 



Figure 7. MIT-Harvey receiver. 




140 


RADIO-CONTROL SYSTEMS 



Figure 8. Philco receiver circuit. 


6 * SECURE SYSTEMS a 

Security of a radio-control system for guided mis¬ 
siles can be obtained by two methods. One method is 
to use secrecy. In the ideal case the enemy would not 
even know that the missile was under control. This 
ideal is probably hopeless of realization, but it might 
be possible to make the system so difficult to analyze 
that the enemy would have to resort to large blocks 
of radiated power over a wide range of frequencies. 
Such a system of jamming would present to the 
enemy serious logistical problems. The second meth¬ 
od is to use a control link in which the receiver an¬ 
tenna is substantially blind to signals not originating 
at or near the control point and to provide sufficient 
strength in the control signal to blast through enemy 
interference. 

Soon after its organization the Division canvassed 
the developmental programs on secure radio-control 
systems under way in the Services. In view of the 
acute shortage of research facilities it seemed inadvis¬ 
able to duplicate any of their projects or to add to 


a See also STR of Division 15, NDRC, for a much more 
complete discussion of jamming and its avoidance. 


their number. Especially was this the case as some of 
them, notably Rex under development for ARL, 
were of great promise. The Division, therefore, under¬ 
took no new projects in so-called jamproof control 
systems. From Section D-3, a predecessor unit of 
NDRC, the Division did, however, inherit one proj¬ 
ect in secure radio-control systems. It is discussed in 
Section 6.4.1. 

Probably the greatest promise of success for an 
invulnerable radio-control system lies in the use of 
coded pulses. If the pulses are made very short, their 
power levels can be very high. The microwave fre¬ 
quency range is extremely suggestive of possibilities: 
for example, the continuous focusing of a receiver 
antenna which would have a narrow field of view on 
the transmitting antenna. Such a receiver would be 
substantially blind to an enemy jamming transmitter 
not collinear with it and its control transmitter. 

6,4,1 Hammond Control System 11 

It has been pointed out (Section 6.4), that one way 
of producing a secure radio link is through secrecy. 
The contractor’s approach to the problem was along 
this route, producing a radio-control link whose 

























































































































































SECURE SYSTEMS 


141 


analysis could be accomplished only with great 
difficulty. 

The system consists of a radio transmitter radiat¬ 
ing power at an average carrier frequency of 84 me. 
This carrier is “wobbled” at a frequency of 440 c 
through a band 20 kc wide. Control is applied by the 


use of two supersonic modulation frequencies, each of 
which in turn is subject to frequency modulation at 
five different frequencies, all of which are at the low 
end of the audio range. The modulating frequencies 
swing about a mean value of 60 and 80 kc. For a 
right-turn signal, the 80-kc ,tone is frequency-modu- 



Figure 9. Philco receiver. 




142 


RADIO-CONTROL SYSTEMS 


lated at 140 c; for a left turn, the same carrier is 
frequency-modulated at 240 c. For an up signal— 
increase in range—the 60-kc tone is frequency- 
modulated at 140 c, and for a down signal the 60-kc 
tone is modulated at 240 kc. When no control is ap¬ 
plied, the 80-kc tone is frequency-modulated at 180 c. 
Thus, the carrier is on continuously, swinging at 440 
c around a mean value of 84 me. It is continuously 
amplitude-modulated at a control tone with a mean 



Figure 10. Comparative selectivity of A, AN/CRW- 
2A; B, MIT-Harvey audio AGC; C , Philco; and D, 
MIT-Harvey conventional AGC. 


value of 80 kc, and this control tone is in turn fre¬ 
quency-modulated at 140, 180, or 240 c. A second 
amplitude-modulating control tone of 60 kc is on 
intermittently and is frequency-modulated at either 
140 or 240 c. 

To produce such a complicated energy array re¬ 
quired a somewhat complicated transmitter circuit. 
Even its block diagram is not reproduced here. The 
decoding of the signal to obtain control operation of 


the missile is equally complicated. It consists of (1) a 
superheterodyne receiver, followed by a detector to 
eliminate the 84-mc carrier; (2) filters to separate the 
60-kc control tone from the 80-kc control tone; 
(3) discriminators to eliminate the low-frequency 
modulation; and (4) a set of filters to energize the 
proper relays from the 140-c or 240-c voltages pro¬ 
duced. An audio AGC is provided which is actuated 
by the voltage following the low-frequency dis¬ 
criminators. 

The purpose of wobbling the carrier is to make 
sure that no single jamming note present on the air 
will beat with the control carrier and produce false 



16 17 18 19 20 21 22 23 24 25 26 27 28 

BATTERY VOLTAGE 


Figure 11 . Comparative sensitivity of A, A/N-CRW- 
2A; B, Philco; and C, MIT-Harvey receivers. 

operation of the relays. The purpose of the super¬ 
sonic amplitude-modulating control tones is to in¬ 
crease the difficulty of analysis. The low-frequency 
modulation provides, among other things, for obtain¬ 
ing more than one control function from a single 
amplitude-modulating control tone. The purpose of 
the audio AGC has already been discussed (Section 
6.3.2). The other purposes of these circuit elements 
are explained in the following quotation from the 
contractor’s final report. 11 

a. Use of Superaudible Modulation. The prime reason for 
using superaudible rates of amplitude modulation is because 










































































































































SECURE SYSTEMS 


143 


usually types of transmitters do not produce such modulations. 
If a CVV or an ICVV signal only is impressed upon the receiver 
in the wave band of the desired signal, then power representa¬ 
tive of the interference will most certainly get through to the 
detector, but the selective nature of the output system of the 
detector will reduce the effect of the undesired signal upon sub¬ 
sequent circuits. Only when the transmitter modulations are 
in the proper frequency range will the effect pass through into 
the subsequent circuits, and even then only when they are 
properly otherwise characterized will the later circuits be 
properly actuated. 

b. Interference Upon Operation by a Standard Transmitter. 
There is of course some considerable virtue in a receiver sys¬ 
tem which cannot be operated by the usual types of trans¬ 
mitters. But the use of such a receiver system does not insure 
that the control exercised by the proper type of transmitting 
systems will not be interfered with by a standard transmitter, 
by causing the final receiver circuits to fail to operate, or to 
operate in an incorrect manner. Three possible sources of dis¬ 
turbance upon operation by the undesired type of transmission 
must be recognized and understood. 

1. Beating Effects. Power which will pass through the receiver 
and develop output voltages can be produced by the conjoint 
action of two transmitters. For example, if the wobbler of the 
radio transmitter of the present equipment, and also the modu¬ 
lation transformer, are not actuated, then the output of the 
receiver unit will be actuated if the transmitter carrier is 
heterodyned by any continuous wave transmitter either 60 or 
80 kcs. different from the proper 84 mcs. carrier. If now the 
wobbler is set into operation, then the beat due to the con¬ 
joint action of the transmitter and the interference will not be 
steady, but will be frequency modulated, over a wide range. 
As a result the root mean square value of the voltage at the 
radio receiver output terminals will be much less than if a con¬ 
tinuous beat note were present, of proper value. The same ef¬ 
fects hold if the transmitter is modulated in such a manner 
that its output is analysable into a plurality of continuous 
waves. The artificial wobble of the transmitter, in short, is 
for the purpose of scattering or distributing the beat frequen¬ 
cies due to the conjoint action of transmitter, modulated by 
signal forming circuit, and the action of continuous wave type 
interferences. The range of wobble is not desired to be so great 
as to vary appreciably the ability of the radio receiver to pro¬ 
duce the desired output in an efficient manner. 

2. Blocking Effects. Interfering energy which is impressed 
upon the detector together with the desired energy can modify 
the ability of the detector to produce the desired amount of 
output. The detector sensitivity for the proper signal is a func¬ 
tion of the ratio of interference to signal. This is not necessarily 
a question of overloading, but in the case of linear detection 
is a natural consequence of the nature of the linear detector 
report. Unless a “square law” detector is used, which would 
bring in other undesirable effects, the loss of detector sensitiv¬ 
ity as the result of the presence of the interference must be 
compensated for by use of automatic signal leveling devices 
subsequent to the detector. 

3. Shocking Effects. Usually the detector circuit is so con¬ 
nected that the voltage impressed upon the detector input 
changes the dc current through a filter coil of the selective 
output system, or changes the dc voltage across a filter con¬ 
denser. If therefore the detector is energized by a continuous 


wave signal alone, and this is abruptly changed in strength as 
by telegraphic keying, then there will be corresponding changes 
in the energy content of the filter device. If the circuits are 
too sharp, oscillations may be set up in the filter in the process 
of changing its energy content, so that voltages of frequencies 
which the filter would transmit may be produced at the radio 
receiver output, even though the proper transmitter to produce 
such voltages in a sustained manner is not operative. 

These three effects of beating, blocking and shocking either 
singly or conjointly limit the ability of the radio receiver cir¬ 
cuit to deliver to the signal reproducing circuit a wave form 
corresponding to the transmitter modulation transformer cur¬ 
rents, in the face of excessive interferences. 

c. Volume Control. When Multiplex transmission is used, as 
required for the present equipment, the selective channels 
must be kept at about proper signal level, or else a signal which 
is sufficiently strong will be able to actuate also the adjacent 
channel. In general, the lower the degree of selectivity between 
adjacent channels, the greater is the need for good volume 
control operation. Proper adjustment of signal level in the 
selective channels may be made by instantaneously operative 
volume control devices, such as chopper type limiters, signal 
amplitude limiters, or by integrating devices such as rectifiers 
operated by the signal to produce dc which in turn controls 
the gain of amplifiers prior to or succeeding the signal channel 
in the receiver chain. The usual Automatic Volume Control 
systems commonly found in broadcast and communication 
equipment produce a rectified and integrated dc control volt¬ 
age in accordance with the amount of total signal impressed 
upon the detector. This arrangement most definitely should 
not be used when interference is expected upon the receiver 
operation, since the gain of the receiver would be diminished 
even when the true desired signal radio field is constant. Even 
with reasonable precautions, the amount of operating signal 
may be decreased by the interference, and very best design 
may require compensatory gain control of the signal system 
in accordance with the ratio of interference to signal. 

In summary, the radio transmitter-receiver combination 
must be coordinated so that (a) standard transmitters in com¬ 
mon use cannot of themselves cause operative functioning of 
the final receiver circuits; (b) standard transmitters cannot 
readily disrupt operation by the proper signal by (1) beating 
effects, (2) blocking effects, or (3) shocking effects; and 
(c) standard transmitters cannot readily control the signal 
level in the final circuits corresponding to the desired signals, 
through improper arrangement of the volume control system. 

The foregoing discussion of invulnerability to jam¬ 
ming is supported at length by mathematical anal¬ 
ysis in the contractor’s final report. No tests were 
reported by him, however, as supporting his assump¬ 
tions and the mathematical analyses developed there¬ 
from. The Division’s consulting engineers at MIT 
(Contract OEMsr-240) made antijamming measure¬ 
ments on the system in the same manner as has 
already been discussed in connection with the MIT- 
Harvey receiver. No significant difference was ob¬ 
served in the performance of the two systems. 



144 


RADIO-CONTROL SYSTEMS 



6 5 EARLY STUDIES 

IN PROPORTIONAL CONTROL 8 

This chapter on radio-control systems b should not 
close without a brief comment on a research program 
carried out by Purdue University at the instigation 
of Section D-3, NDRC. The purpose of this program 
was to develop a multichannel proportional control 
link. The scheme developed was to modulate a carrier 
with one audio tone for each control channel. The 
frequency of each audio tone was varied in propor¬ 
tion to the control desired. In the receiver sufficiently 
broad-pass filters separated the channels and drove a 
discriminator stage for each (Figure 12). The result 
was a d-c voltage for each channel proportional to the 
controlling signal. 

The audio tones were provided from resistance- 
capacitance oscillators consisting of two twin triodes 
in a phase-inverter circuit. One grid of each twin 
triode was grounded. The other grid was driven from 
an appropriately tuned phase-shifting circuit across 
the plates of the second tube. The tubes operated, 
then, in push-pull output with single-ended input. 
This type of oscillator gave good stability and wave- 

b See Chapter 4 for further discussion of methods of obtain¬ 
ing proportional control. 


form, with economy of tubes over conventional oscil¬ 
lator design. The outputs of the oscillators for two 
channels were combined in a mixer stage driven by 
cathode followers. The mixed signals amplitude- 
modulated the carrier in the usual manner. 



Figure 13. Discriminator and motor control of Pur¬ 
due receiver. 


The development of a discriminator for the a-f 
range presented some new problems. The final design 
is shown in Figure 13. Twin diode discrimination is 
used as in conventional r-f practice. In the r-f appli¬ 
cation, however, the input transformer is tuned so 
that the voltage across the anodes of the diodes is at 
90 degrees with the primary voltage. In the a-f appli- 






















































































EARLY STUDIES IN PROPORTIONAL CONTROL 


145 


cation an iron-core transformer is used whose second¬ 
ary voltage is substantially in phase with the primary 
voltage. The tuned circuit LiC 2 is resonant at the 
frequency corresponding to a zero control signal. It is 
excited at 90 degrees from the primary through the 
phase-shifting circuit consisting of Ci, C 3 , and the 
tuned loop. Li is made variable, and its value is de¬ 
termined by the servomotor. Thus, if the control- 
channel frequency is raised, a voltage will pick up the 
relay R F, energizing the servomotor which readjusts 
Li until balance is restored. 

It should be noted that this system is not “fail¬ 
safe.” The plate current of the d-c pentode stage is 


adjusted midway between the pickup and drop-out 
values for the relay. If radio transmission is lost, the 
servomotor will continue to run to the limit of its 
travel in whichever direction the relay last happened 
to be. 

Considerable difficulty was experienced with hunt¬ 
ing of the servomotor. Although, through an ingen¬ 
ious vernier on L h the contractor solved that insta¬ 
bility, his success is no indication that hunting would 
not have appeared if the system had been applied to 
a missile. Hunting is a servo-system problem. If it is 
attacked piecemeal, success will result only by 
chance. 



Chapter 7 

SERVOMECHANISMS 


71 INTRODUCTION 

S ervomechanisms are control devices to impose 
an output which must vary in time according to a 
varying input signal. The power for this output must 
not come from the input but must be supplied locally. 
The operation of the device depends upon the error 
signal obtained by comparing the existing output 
with the existing input and driving the device in the 
sense to minimize this error signal. Servomechanisms 
are the mechanical analogue of feedback amplifiers 
and are subject to similar instabilities. Much of the 
recent progress in the theory of servomechanisms is 
due to the recognition of this analogy. 

Much progress was made during the war in the 
development of servo theory and in its application to 
fire-control problems. Section D-2 of NDRC, its suc¬ 
cessor Division 7 (Fire Control), and the Applied 
Mathematics Panel all promoted investigations in 
this field. Of the several reports written as a result of 
these activities, the reader is particularly referred to 
one by MacColl, 1 from which the definition of a servo¬ 
mechanism given above is paraphrased. A homing 
missile is a particular type of servomechanism. Its 
direction of flight is controlled by the error signal 
between its present direction of flight and the direc¬ 
tion of the target. This again is a feedback process 
liable to instabilities. Servomechanisms were devel¬ 
oped for the control of the various homing missiles of 
Division 5. (If the human operator is included, the 
control systems of Azon, Razon, and Roc are also 
servo systems.) These have already been discussed; 
one project, however, deserves special discussion and 
has been reserved for this chapter. 

The Servomechanisms Laboratory of the Massa¬ 
chusetts Institute of Technology had played an im¬ 
portant role in the development of servo theory and 
had been successful in applying it to the design of 
fire-control equipment. Shortly after Division 5 was 
organized, the laboratory was persuaded to accept a 
contract to apply the same principles to the design of 
a stabilizing and control system for the glide bomb 
known as Pelican. Before the work was completed, 
Service interest shifted to Bat, and the design 
of the control system was altered to fit this missile. 
A satisfactory control system was developed, and 


a production design was in progress as the war 
ended. 

Pelican and Bat will soon become obsolete with 
the advent of powered missiles of long range, 
but the philosophy of design of the control system 
and the dynamic methods of testing will, with proper 
modifications, be applicable to a wide variety of 
future missiles. Increased speed and cost of future 
missiles will lend additional importance to the meth¬ 
ods of design and of simulative testing. The two re¬ 
ports submitted by the laboratory 2 ’ 3 in their entirety 
are strongly commended to the attention of all 
guided-missile engineers. The abstracts to be pre¬ 
sented here are accordingly brief. 

72 CONTROL SYSTEM 

FOR PELICAN (RHB) MISSILE 

721 Principles of Construction 

The first report 2 from the Servomechanisms Labor¬ 
atory at MIT covers the work from May 1, 1943, to 
March 15, 1945. Because homing control was re¬ 
quired, a particular airframe (Figure 1) had been 
designed by the National Bureau of Standards to 
give as nearly constant an angle of attack as possible. 
For this reason the conventional airplane arrange¬ 
ment of rudders, elevators, and ailerons was aban¬ 
doned in favor of full-span trailing-edge flaps called 
elevons, which performed the functions of all three 
types of conventional control surfaces. Differential 
action of the elevons produced the effect of ailerons in 
roll control; cooperative action changed the lift of the 
wing and thus regulated the angle of glide. Turning 
was accomplished by the horizontal component of 
lift when the missile was put into a bank by the 
differential action of the elevons. All this has been 
described in detail in Chapter 1. It is set forth briefly 
in the report from the Servomechanisms Laboratory 
as it governs the design of the servo-control system. 
Advantages and disadvantages of this type of aero¬ 
dynamic control are also discussed. 

In place of the rate gyros used in the NBS system, 
free vertical gyros were selected. The first model 
utilized a vertical gyro manufactured by Minneapolis- 
Honey well. It was altered to employ rotatable trans- 


146 


CONTROL SYSTEM FOR PELICAN (RHB) MISSILE 


147 


formers to indicate gyro angles in pitch and roll. The 
stators of these rotatable transformers were in turn 
rotated by trim motors to permit changes in the re¬ 
quired angles of glide and of bank. In a second model 
a Sperry Mark IV bank and climb control unit was 
used which included rotary air valves to measure the 
same two angles. The second model used a smaller 
gyro and required a minimum of electronic equip¬ 
ment; the comparison between gyro position and 
elevon position was made mechanically through a 
linkage to the air valves, with differential introduc¬ 
tion of trim corrections into this linkage (Figure 2). 

In both systems an error of roll position or of glide 
angle called for elevon action of an amount to set up 
restoring forces proportional to the magnitude of the 
error. This was true whether the error was measured 
from the original preset flight orders or from the 
altered orders introduced by trim motors responsive 
to the instructions of the radar receiver. In the pitch 
control a small permanent-magnet generator, geared 
to the pitch trim motor, holds the rate of elevon mo¬ 
tion proportional to the radar error signal. 

After an initial attempt to drive the two elevons 
independently, a linkage system was developed to 
control the two elevons from a single motor through 


four magnetically actuated clutches. The motor was 
nonreversible, and pairs of clutches connected it in 
either sense to each of two points in the linkage. Mo¬ 
tion at one point in the linkage drove the two elevons 
in opposing directions, while motion at another point 
drove them together. This device secured rapid re¬ 
sponse with high motor efficiency. The relatively 
short life of the clutches (20 hours) was entirely 
adequate for the testing and use of the missile. 

7 2,2 Flight Test Table 

The portion of the report of most continuing sig¬ 
nificance is a description of a laboratory flight table 
(Figure 3) for the simulative testing of the control 
system and of its components. Construction of this 
device consumed considerable time but paid valuable 
dividends. The table carried the radar receiver and 
gyro and pointed toward a radar target set up a short 
distance in front of it. Only those portions of the sys¬ 
tem sensitive to the motion of the glider had to be on 
the table. The servomotor itself was set up on a 
near-by bench, and its output was loaded to repre¬ 
sent the elevon wind loads. The position of the ele¬ 
vons in turn controlled the motions of the table. Rate 



Figure 1. Radar-homing glide bomb with MIT stabilizing and servomechanism. 



148 


SERVOMECHANISMS 


of roll of the table was governed by the differential 
displacements of the elevons. Roll orientation of the 
table determined its rate of rotation about a vertical 
axis. Pitch was simulated by raising and lowering the 
radar target. The electric, hydraulic, and mechanical 
components used to accomplish this are described in 
the report. It is sufficient here to state that the oper¬ 
ation of the table approximated the dynamics of the 
missile flying at large ranges from the target. 

Tests using this table gave the quantitative data 
needed to determine the dynamical characteristics of 
the overall control system at a fraction of the cost in 
money and in time that would have been needed to 
obtain the same data by flight tests. There is the 
further advantage that in laboratory tests all the 
variables are under the control of the investigator, 
whereas in flight tests this is far from true. Some of 
the flight responses which were studied were (1) the 
accuracy with which the missile would hold a course; 
(2) the damping of the missile to suddenly applied 
roll moments; (3) the nature of response to changes 
in flight direction. In general the control constants 



determined from the flight table were found to oper¬ 
ate well with the missile in actual flight. 

In addition to its use in establishing design para¬ 
meters for the MIT servomechanism, the table was 
used to examine other stabilization systems for 
Pelican. The control system developed at the Na¬ 
tional Bureau of Standards (see Chapter 1) was also 
tested on it to find optimum adjustments. 

7 2 3 Preliminary Flight Tests 

Detailed results are given of a number of flight 
tests with each of the two systems described. These 
tests were made against a radar beacon target sus¬ 
pended from a barrage balloon. The tests and the 
parachute recovery device to preserve apparatus and 
flight records are summarized in Chapter 1 and 
Chapter 8. Four units were tested with the Minne- 
apolis-Honeywell gyros and rotatable-transformer 
pick-offs. The first rolled excessively and tumbled the 
gyro, then went into a steep dive until the parachute 
recovery device operated. The second passed about 






















































































































CONTROL SYSTEM FOR BAT (SRB) MISSILE 


149 



90 ft to the left of the target and about 90 ft high; a 
cross wind of approximately 20 mph was blowing 
from the right. The third test unit flew an oscillating 
course toward the target but the parachute opened 
prematurely before reaching the target, and the mag¬ 
nitude of the miss could not be determined. The 
fourth unit lost its radar signal during an instability 
early in the flight but later flew a smooth course 
without radar. 

Four units were tested with the Sperry gyro and 
the pneumatic take-offs. The first test failed because 
of premature functioning of the recovery parachute. 
The second unit homed on the beacon and cut the 
cable of the beacon 18 ft above the target. In the 
remaining tests the beacon was mounted 12 ft above 
the ground. The third unit flew well: its nose passed 
15 ft to the left of the target and 8 ft above it. The 
fourth unit passed 18 ft to the right of the beacon and 
17 ft above it. 

™ CONTROL SYSTEM 

FOR BAT (SRB) MISSILE 

As a result of the work described in the contractor’s 
preliminary report, 2 it was decided to develop this 
type of control for use in Bat. As has been detailed 
elsewhere, Bat is somewhat larger than Pelican, car¬ 
ries a heavier bomb, and has a send-receive radar 
rather than the simple receiver used in Pelican. The 


final report 3 from the Servomechanisms Laboratory 
is a detailed description of the system developed and 
of its components. Some of the material contained in 
the first report is repeated where it is necessary to 
elaborate it, but the second report should not be read 
without the first. Some thirty pages of analysis are 
given which apply to this problem the analytical 
methods referred to in Section 7.1. The purpose and 
scope of this analysis is best described by a quotation 
from the report: “Experience with a variety of con¬ 
trol problems has proven that the system analysis 
must yield much more than an answer as to whether 
the system is stable or unstable. The analysis must 
lead to a design with a determinable and satisfactory 
margin of stability, must determine and control such 
factors as the natural frequencies of the complete 
system and its damping characteristics, and must 
establish the system sensitivity factors.” A summary 
is given of servo theory, 4 so succinct that it is repro¬ 
duced here. 

7,3,1 Dynamical Analysis 

of System Performance 

Frequency Response of an Automatic 
Control System 

“Most automatic control problems can be reduced 
to a simple block diagram as illustrated by Figure 4. 






































150 


SERVOMECHANISMS 


The problem illustrated by this diagram is that of 
making an output shaft whose motion is designated 
by 6 0 follow an input shaft whose motion is designated 
by Si. The servomechanism approach to the solution 
of this problem is to compare the output 6 a with the 
input 6i and use the difference between these quan¬ 
tities to so operate the controller and servomotor that 
the difference between the output and input is made 
zero or minimized. The difference between the output 
and the input is termed the ‘error’ and denoted by e. 



Figure 4. Block diagram of automatic control system. 


variety of applications, the best system response is 
underdamped, but with a high damping factor, as 
illustrated by the response curve so marked. The 
length of time that the output requires to reach a new 
position is a function of the damping and frequency 
characteristics of the oscillation. An estimate of the 
length of time required for a transient to disappear 
is given by a knowledge of these two factors. 

“The transient approach to the problem is satisfac¬ 
tory if the response characteristics as calculated or 
measured are satisfactory or will be satisfactory with 
only minor readjustment of system parameters. In 
general, however, it is very difficult to design or re¬ 
design a system from the transient standpoint. 

“The second method of approach to the design and 
analysis problem is to study the response of the sys¬ 
tem to a sinusoidal input. In this approach 0* is made 


“Two general methods of studying the response of 
such a system are in use: the first method studies the 
transient response of the system, and the second 
method studies the sinusoidal response of the system. 
In the study of the transient response, the procedure 
is to displace the input 0* in accordance with some 
transient test function and measure or calculate the 
response of 0 O . A common test function is a step func¬ 
tion as illustrated by Figure 5. The output can follow 
the input in a number of ways, depending upon the 

$i<t) 



Figure 6. Responses of automatic control system to 
step displacement. 


i 


o ------ 

t 

Figure 5. Input step displacement. 

characteristics of the amplifier-controller-servomotor 
combination. Certain types of output response to a 
step input are indicated in Figure 6. If the system 
parameters are mischosen, the system may be un¬ 
stable, and output will have a continuing oscillation 
illustrated by the unstable response. For another se¬ 
lection of parameters the output may approach its 
required position in a very slow fashion. This response 
is called overdamped and for many applications is 
as unsatisfactory as the unstable response. For a 


a sinusoidal function of constant magnitude and vary¬ 
ing frequency, and the steady-state performance of 
the output is calculated or measured. 

“The response of the system to this type of input 
is prescribed by determining as a function of fre¬ 
quency, first, the ratio of the amplitudes of the out¬ 
put motion and, second, the relative phase between 
the output motion and the input motion. Mathe¬ 
matically speaking, if 0<ft) is the input and 0 o (t) is 
the output of a servomechanism, then, if 

Bi(t) = A sin cot (1) 

and A is small, it will always be true that 

d 0 (t) = B sin (cot + <t>) (2) 

“In the above equations A is known as the ampli¬ 
tude of the input, B is the amplitude of the output, 
co is the angular frequency (equal to 2irf where / is 




















CONTROL SYSTEM FOR BAT (SRB) MISSILE 


151 


in c) of motion of Oft) and Oft), and <£ is the relative 
phase angle between Oft) and Oft). 

“The amplitude ratio B/A and the phase angle </> 2 
when determined as functions of angular frequency co 
comprise the frequency-response characteristic of the 
system. 

“As shown above, the relationship between 0 o {t) 
and Oft) when 6ft) is a sinusoidal function of time is 
specified by a magnitude B/A and an angle <t>. These 
two quantities can be considered as the defining prop¬ 
erties of a vector whose amplitude is B/A and whose 
phase is <t>. Thus it is frequently stated that a vector 
relationship exists between 6ft) and 0 o (t) when Oft) 
varies sinusoidally with time. This vector relation¬ 
ship is represented symbolically by (0 o /0i)(jco) in 
which j(= \/—l) itself is generally thought of as a 
vector and emphasizes the vector properties of the 
ratio. The vector ratio ( Oo/Ofiju ), is characterized 
by an amplitude | {Oo/Ofijco) | and a phase, arc 

“The preceding development is briefly summarized 
as follows: 

If in an automatic control system 

Oft) = A sin oit (1) 

then 

Oft) = B sin {(^t + <t>) (2) 

By definition, the frequency response is denoted by 
(i 0 o /0i)(joi ). The amplitude response is given by 


and the phase response is represented by 



“The amplitude and phase response curves of a 
typical system are illustrated by Figure 7. 

“The curves of Figure 7 may be obtained (1) by cal¬ 
culation, if the constants of an actual or proposed 
design are available, or (2) by measurement, if the 
servomechanism itself is available. The frequency- 
response characteristic is measured by moving the 
input sinusoidally at a fixed amplitude but at various 
frequencies. At each frequency the amplitude of the 
output and its phase relation to the input motion is 
measured. The ratio of the amplitudes, plotted for 
each frequency, yields the first of the curves of Fig¬ 
ure 7. The phase difference between the input and 
output, plotted for various frequencies, gives the 
second of the curves of Figure 7. The curves of Figure 


7 are readily calculated if the constants of the system 
and the differential equation relating the output to 
the input is known. The calculation is effected by re¬ 
placing the operator p by the frequency operator jco 
and applying conventional vector arithmetic. 

“The frequency response of a servomechanism may 
be closely correlated with its transient response. Im¬ 
portant natural frequencies in the transient response 
are indicated by peaks in the amplitude-response 
curve. The magnitudes of the peaks of the amplitude 
response are measures of the relative damping of the 
natural frequencies of the transient response. The 
frequency band over which the amplitude response 
has a substantially constant magnitude is a measure 
of the speed of response to transients, since a high 
natural frequency (and, therefore, a high speed of 
response) is linked with a high resonant frequency in 



Figure 7. Frequency response characteristics. 


the amplitude response. When, as indicated above, 
the frequency-response characteristic is correlated 
with the transient-response characteristic, the former 
becomes a powerful means of analysis. 

“The correlation between the sinusoidal and tran¬ 
sient characteristics is illustrated by comparing the 
transient and frequency response of a simple system 
representable by a second-order differential equation. 

“If the input Oft) is a step displacement and the 
output Oft) is determined for various values of the 
damping ratio, the set of transient responses illus¬ 
trated by Figure 8 results. If Oft) is made a sinusoidal 
function, 

Oft) = A sin oit (1) 

and the amplitude response of the output 0 o (t) is de¬ 
termined, the set of curves of Figure 7 are obtained. 
Comparison of the transient and frequency responses 
reveals a number of points of correspondence: (1) the 
frequency at which the transient response oscillates 










152 


SERVOMECHANISMS 


(the natural frequency) is approximately the same as 
the frequency at which the amplitude response has a 
peak (the resonant frequency); (2) as the damping 
ratio is reduced, the transient response becomes more 
oscillatory, and the peak in the amplitude response 
is magnified; (3) when the damping ratio is made 
larger than unity, the transient response becomes 
sluggish, and the amplitude response falls off rapidly 
without a peak. 

“It is frequently convenient to combine the infor¬ 
mation contained in the amplitude-response curve 
and in the phase-response curve of a system by a 
single graph. This can be accomplished by remember¬ 
ing that the amplitude ratio and the phase angle are 



0 2 4 6 8 


Figure 8. Responses of second-order system to step 
displacement. 

quantities defining a vector which relates the output 
and input. As the frequency of the input is varied, 
this vector varies in phase and magnitude, and the 
amplitude- and phase-response curves present the in¬ 
formation on the manner in which these two proper¬ 
ties of the vector change with frequency. The same 
information can be provided in an alternative way 
by plotting on polar coordinate paper the path fol¬ 
lowed by the tip of the vector as the frequency varies 
over the band of interest. If various points of the 
curve are labeled with the frequency to which they 
correspond, one curve will supply the information 
contained in two curves when the amplitude and 
phase response are plotted separately. This graphical 
presentation is known as the locus of the frequency- 
response characteristic. It is illustrated by Figure 9 
which was drawn for the same second-order system 


to which the curves of Figures 7 and 8 apply. The 
principal advantage of presenting information in this 
fashion lies in the means it provides for visualizing 
the frequency response and in the fact that it em¬ 
phasizes the very important relationship that always 
exists between the amplitude and phase responses of 
a system. 

Transfer Function of an Automatic 
Control System 

“It is clear from the block diagram of Figure 4 that 
a single function completely defines the performance 
of a servomechanism in which the feedback link con¬ 
tains no frequency-dependent elements. The defining 
function is the relationship between the servo output 
S 0 (t), and the error e(t). If this relation is known in 


IMAGINARY AXIS 



Figure 9. Complex plane plot of sinusoidal responses. 


operational, sinusoidal, or time-response form, the 
performance of the system is completely defined for 
all conditions that may be imposed upon it. This 
function, relating the system output of the closed- 
cycle system to its error, has been termed the transfer 
function of the system. When the sinusoidal or fre¬ 
quency-response form of the transfer function is 
studied, it becomes a powerful analysis and synthesis 
tool. The transfer function may be derived from a 
known frequency-response characteristic of the servo 
system, it may be measured directly, or, if system 
constants are known, it may be calculated. The trans¬ 
fer function of the system can be written in the form 
(6 0 /E)(j a>). This function is a vector quantity with 
an amplitude and a phase characteristic just as the 
ratio ( 0 o /di)(jo3 ) is a vector quantity. 

“The transfer function of a closed-cycle system is 
always the product of two parts, one that is invariant 
with frequency and a second that is frequency de¬ 
pendent. The fact that these two components exist is 





















CONTROL SYSTEM FOR BAT (SRB) MISSILE 


153 


emphasized by writing the transfer function in the 
following form: 

f-0'«) = KG(Jo>) (5) 


“The term K represents that part of the transfer 
function which is invariant with frequency. This por¬ 
tion is known as the gain or the sensitivity factor and 
is a function of amplifier gain, etc. The second part 
of the transfer function is denoted by G(jco) and rep¬ 
resents the portion of the transfer function which 
changes with frequency. 

“The frequency response of the servomechanism is 
related to the transfer function by the following vec¬ 
tor equation: 



KG(ju) 

1 -f KG(ja>) 


( 6 ) 


“The transfer function of a system is calculated by 
straightforward circuit analysis techniques, or by de¬ 
termining the differential equation relating the out¬ 
put to the error and replacing d/dt by jw. 


Figure 10. Suppose it is desired to calculate the mag¬ 
nitude and phase of (0 o /0i)(j co c ) for a particular fre¬ 
quency co c with only the transfer locus available. The 
transfer function at co c , KG(joj c ), is represented by 
the vector oc while the vector ac represents the term 
[1 + KG(ju c )], since the point a is located at 



Figure 10. Transfer-function locus. 


The Transfer-Function Locus 

“The transfer function can be studied by means of 
its frequency characteristics just as other functions 
have been so studied. The amplitude- and phase- 
response curves of the transfer function can be drawn; 
these curves completely define the characteristics of 
the transfer function of the servomechanism and 
therefore completely define the system itself. Just as 
the phase- and amplitude-response curves were com¬ 
bined into a single polar plot with frequency as a 
parameter, so can the frequency- and phase-response 
curves of the transfer function be combined. This 
parametric polar plot of the transfer function has 
been called the transfer-function locus, or simply the 
transfer locus of the servomechanism. The transfer 
locus completely defines the characteristics of the 
servo system. A study of its nature provides a very 
effective method for the synthesis of servomecha¬ 
nisms intended for particular applications and a use¬ 
ful general guide to the adjustment of servomecha¬ 
nism parameters to secure optimum performance. 

“If a parametric frequency plot of the transfer func¬ 
tion KG(ju) is available, the magnitude and phase 
of the function ( 0 o /0i)(jo >) may be found by graphical 
calculation. The function ( 0 o /0i)(ju ) has been related 
to the transfer function by equation (6). A plot of 
a typical transfer function KG(jco) is illustrated by 


(—1 + j0). Therefore, the magnitude of (0 o /0i)(ju c ) 
is given by 



\KG<M 1 = |oc| 

\l+KG(jw c )\ | oc | 


while the phase of ( 6 0 /$i)(ja ) is given by 


[ Jovi] = 


arc (oca) 


(7) 

(8) 


The angle, arc (oca), is negative. Thus the magnitude 
of ( do/di) (jo) c ) is equal to the ratio of the magnitudes 
of the vectors oc and ac and the phase of (0 o /0i)(jo) c ) 
is equal to the angle between these two vectors. 

“Equations (7) and (8) permit ready visualization or 
calculation of the magnitude and phase of (0 o /0i)(joi). 
At small frequencies, such as co b (see Figure 10), both 
vectors oc and ac are large and approximately equal, 
and their ratio is approximately unity. The angle be¬ 
tween the two vectors, the phase of ( 0 o /0i)(ju), is 
small at this frequency. As the frequency increases, 
the angle between the two vectors increases, and their 
lengths become smaller, so that differences in their 
lengths cause the ratio | oc | /1 ac | to depart from 
unity. Whether the ratio | oc | /1 ac | [the magnitude 
of (0 o /0i)(ju)] increases or decreases as the frequency 
increases depends upon the shape of the curve rela¬ 
tive to the origin and the point (— 1 + jO) . A con- 









154 


SERVOMECHANISMS 


tinuation of this reasoning for the remainder of the 
frequency range permits the general shape of the 
phase and magnitude of the servo output to be com¬ 
pletely determined. If desirable, the phase and mag¬ 
nitude curves can be determined from the transfer 
locus with accuracy and ease by the use of a protrac¬ 
tor and divider to measure the angles and lengths 
directly from the graph. 

“The general nature of the amplitude response ma}^ 
be obtained also by drawing in the complex plane 
curves of constant | (6 0 /di)(jo })\. These curves are 
circles and are shown in the upper part of Figure 11. 
If a servo could have a transfer-function locus that 
lay along one of these circles, the system amplitude 
response would be independent of frequency and 
equal to the value of M for which the circle was 


imaginary axis 



Figure 11 . Determination of servo sensitivity. 


drawn. The center and radius of each circle is de¬ 
pendent upon M, and the relation is given in Figure 
11. The frequency at which the transfer-function 
locus crosses one of these circles is the frequency at 
which the magnitude of | (O o /0i)(j<a)\ is equal to the 
value of M for which that circle is drawn. If the locus 
is tangent to a circle it indicates that a maximum or 
minimum in | (0 o /0*) (jco) | occurs at that frequency. 
Thus the amplitude response of the system whose 
transfer function is plotted in Figure 11 has a peak 
of 1.5 at a frequency co = 2, if k = 1.0. 

Absolute Stability Criterion 

“The primary requirement that almost every servo¬ 
mechanism must satisfy is that of stability. Although 


a servo system must be more than barely stable to be 
satisfactory, a stability criterion of one type or an¬ 
other is generally the first test applied to proposed 
servomechanism design. Several criteria exist; how¬ 
ever, the one described here was developed primarily 
for application to feedback amplifiers. 

“It is at once apparent to those familiar with feed¬ 
back amplifier theory that the transfer locus of a 
servomechanism is analogous to the Nyquist diagram 
of a feedback amplifier. The term Nyquist diagram 
has been given to this type of plot for a feedback 
amplifier because of a very useful criterion developed 
by Nyquist for determining the stability of a feed¬ 
back amplifier. This criterion may be applied equally 
well to the transfer locus in order to determine from 


IMAGINARY AXIS 



Figure 12. Transfer loci of closed form. 


its shape and position whether or not the servo¬ 
mechanism for which it is drawn is stable. To apply 
the Nyquist stability criterion to servomechanisms 
the following procedure is employed: (1) the transfer 
locus KG(ju), plotted in polar form, is drawn for all 
frequencies from zero to infinity; (2) the conjugate 
of the transfer locus is drawn (the conjugate of a 
curve is the mirror image of the original curve about 
the real axis); (3) if the curves so formed enclose the 
point (-1 +,;0), the system is unstable; if the 
curves do not enclose this point, the system is stable. 
The application of this criterion is illustrated by 
Figure 12. 

“The above criterion applies to curves of closed 
form; that is, it applies to transfer loci of such char¬ 
acter that the loci and their conjugates join at zero 
and at infinite frequency. Actually the transfer loci 
of most servomechanisms are of the open form, and 









CONTROL SYSTEM FOR BAT (SRfi) MISSILE 


155 


some extension is required in order to apply the 
stability criteria to these forms of transfer loci. The 
open form of the transfer locus can be changed into 
the closed form by connecting the curve and its con¬ 
jugate at the zero frequency point by means of a 
circle of infinite radius. The connection should always 
be made in such a way that no phase discontinuity 
occurs along the path of the curve. This is illustrated 
by Figure 13, in which are plotted the transfer loci 
of two common types of servomechanisms, both of 
which are stable. 

Transfer Loci for Various Types of 
Steady-State Performance 

“The performance of a servomechanism under 
steady-state conditions is always of great importance. 
If the servomechanism is primarily a positional de¬ 
vice, it is desirable that the servomechanism take up 
various positions without requiring an error to main¬ 
tain it in that position. Similarly it is frequently nec¬ 
essary for the servomechanism to follow an input of 
constant velocity, that is, one in which 6i{t) = kt. 
In this case it is desirable for the servomechanism to 
follow various velocities as required without the ne¬ 
cessity of a system error to maintain that velocity. 
Servomechanisms that satisfy the first condition fre¬ 
quently are termed zero-displacement-error servo¬ 
mechanisms, and those that meet the second condi¬ 
tion are called zero-velocity-error servomechanisms. 
It can be readily shown that if the transfer locus of a 
servomechanism approaches infinity along the nega¬ 
tive imaginary axis, the servo will have zero displace¬ 
ment error. Similarly if the transfer locus of a servo¬ 
mechanism approaches infinity along the negative 
real axis that servo will have zero velocity error. The 
zero-velocity-error servo will, of course, have zero 
displacement error also. In Figure 13 is illustrated 
the transfer loci of a zero-velocity-error system and 
of a zero-displacement-error system. This concept 
can be extended to servo systems that will follow a 
constant input acceleration without steady-state er¬ 
ror, and so forth. 

Determination of the System Sensitivity 

“A very simple procedure exists for determining the 
system sensitivity K permitted by a prescribed maxi¬ 
mum value of | ( 0o/0i ) (jco) |. This procedure makes use 
of the fact that variations in the sensitivity K are 
equivalent to changes in scale of the plot of KG(joj). 


Instead of plotting KG(j co), the function G(ju>) only 
is plotted. If the scale of this plot of G(ju) is correctly 
altered, then this plot will represent KG{ju>), and the 
factor by which the scale must be altered is equal to 
K. The factor by which the scale of the plot of G(joi) 
must be altered to transform it to a plot of KG(ja ;) 
is determined by the maximum permissible value of 
| (d 0 /di)(jo:) |. The procedure is as follows: 

“It has been shown that if the plot KG(ju) is tan¬ 
gent to a circle whose center i& M 2 /M 2 — 1 on the 
negative real axis, and whose radius is M/(M 2 — 1), 
that the function | (0 o /0r)(jco) | corresponding to this 
KG(ju) will have a maximum value of M. Now a 
circle whose center is at M 2 /M 2 — 1 on the nega¬ 
tive real axis and whose radius is M/(M 2 — 1) 
will have an intercept on the real axis equal to 


IMAGINARY AXIS 



Figure 13. Transfer loci of open form. 


M/(M +1), and the following ratio will be main¬ 
tained : 

Center of circle _ M , . 

Intercept on real axis M — 1 

“The scale factor K can be readily found if a circle 
can be located on the plot of G{joi) that is tangent to 
the locus G{ju>) and whose center on the negative 
real axis is M/(M — 1) times its real-axis intercept. 
The location of such a circle can be found by a cut- 
and-try process with a pair of dividers and is the work 
of but a few moments. Suppose such a circle has a 
center at A, on the G(ju) plane. But it has been 
shown that if the scale were correctly chosen in order 
that the locus be a plot of KG(j co), the center of the 
circle would be at M 2 /(M 2 — 1). Therefore the scale 
must be changed by the factor M 2 /A(M 2 — 1), and 
the sensitivity K M is equal to M 2 /A(M 2 — 1). The 
value of Km is the system sensitivity that will provide 
a maximum value of | (0 o /0;)O' w ) | equal to M. 







156 


SERVOMECHANISMS 


“The above procedure of plotting only G(jw) and 
then determining the sensitivity K is easy to use and 
permits greater freedom in the study of transfer loci 
since it essentially nondimensionalizes the plots as 
far as the sensitivity factor is concerned.” 


7 3 2 Production System 

Later chapters of the report 3 cover system descrip¬ 
tion, component description, test equipment, flight 
tests, and a brief discussion of the production model 
engineered by the Bell Telephone Laboratories in 
close liaison with the Servomechanisms Laboratory. 
The system is essentially that described in the earlier 
report. It uses the Sperry Mark IV bank and climb 
control unit with pneumatic take-off of bank and 
glide angles. Ten units were prepared for flight tests 
—three against corner reflectors at Warren Grove, 
N. J., and the remaining seven against a ship target 
in Pamlico Sound, N. C. In the tests against the ship, 


comparison was to be made with an equal number of 
Bats equipped with the system developed at the 
National Bureau of Standards. 

In the Warren Grove tests the first unit hit the 
ground 60 ft in front of the corner reflector. The 
second unit flew just over the reflector and landed 
approximately 70 ft behind. The third unit mal¬ 
functioned and dove into the ground immediately 
after launching. The tests against the ship target 
were more encouraging. The target ship was 260 ft 
long and 45 ft wide, and had a wooden lattice built 
upon the deck to a height of 30 ft above the water 
line. Six missiles were dropped. There were two 
direct hits, one 20 ft aft of midships, 2 ft above the 
water line, the other 45 ft forward of midships exactly 
on the water line. Three skip hits struck the water 15 
to 50 ft short and skipped into some portion of the 
ship or its superstructure. There was one miss which 
passed over the bow and hit the water 100 ft beyond. 
Units with the NBS system had one hit on the ship, 
two on the top of the lattice, and three misses. 



Chapter 8 

INSTRUMENTATION 


11 INTRODUCTION 

-accurate and minute measurement seems to the 
non-scientific imagination a less lofty and digni¬ 
fied work than looking for something new. But nearly 
all the grandest discoveries of science have been but 
the reward of accurate measurement and patient, 
long-continued labor in the minute sifting of numer¬ 
ical results.” These sentences are from an address by 
William Thomson, Lord Kelvin, on the occasion of 
his installation as President of the Royal Society of 
Edinburgh. 1 In a more pungent, if somewhat less ele¬ 
gant, speech Reichsmarschal Goering pointed out the 
need for scientific and quantitative observation spe¬ 
cifically in connection with the development of 
German guided missiles. 2 

Mature work can go forward neither in the field of 
scientific research nor in that of engineering develop¬ 
ment without rigorous quantitative thinking. The 
requirement of quantitative thinking, however, im¬ 
plies the existence of quantitative data. It is par¬ 
ticularly true in the field of guided missiles that ob¬ 
servation by trained individuals is wholly inadequate. 
Different observers differently located will necessari¬ 
ly have different points of view. To each of them the 
apparent performance of the missile may well be 
wholly different. The analysis of the performance of 
the missile made from the reported observation of 
several witnesses s then a matter of discussion be¬ 
tween the witnesses and the chief of the experiment. 
Always there is a danger that clashes will arise which 
are resolved on the basis of the skill of the observers 
in the field of debate rather than on the technical 
merits of the case. Where the military are involved 
there is a further complication of rank. In a discus¬ 
sion with juniors it is almost impossible for a senior 
officer to be in error. 

For these reasons it is vital that measuring and 
recording instruments be provided which will pre¬ 
serve objective and quantitative data as to the per¬ 
formance of the missile during each experiment. 
Without them the program easily degenerates into 
Fourth-of-July play. It is significant that the German 
developmental program on the V-l missile did not 
reach material success until Hilda Frisch, the famous 
glider pilot, volunteered, or was caused to volunteer, 


to ride a buzz-bomb and take records of its perform¬ 
ance during flight. Fraulein Frisch made several 
flights in buzz-bombs over the Baltic Sea, and the 
data which she took were of unique value to the 
scientists and engineers engaged in the development 
of the weapon. 

82 GENERAL 

821 The Problem 

In order to analyze the performance of a guided 
missile it is necessary and sufficient to measure its 
relative motion with respect to the target, and to 
measure the control applied and the missile’s re¬ 
sponse thereto. 

This is a simple statement of an exceedingly com¬ 
plicated problem. The study of the relative motion 
between the missile and the target involves the posi¬ 
tion, the velocity, and the acceleration of each. The 
study of the control involves the determination of 
what controls should have been given in the particu¬ 
lar system of guidance under consideration, what 
actual control was applied, and, in the case of a 
discrepancy, the measurement of the performance of 
each link in the closed loop of the servo system to 
establish beyond doubt the cause of the malfaction. 
Furthermore, the complete measurement cannot be 
made until the response of the missile to the control 
applied has been determined. 

As regards the missile, the problem involves meas¬ 
urements of position, velocity, and acceleration in six 
degrees of freedom. These can be defined, for exam¬ 
ple, as location of the center of gravity of the missile 
in the range, azimuth, and altitude directions, and as 
rotation of the missile’s structure about the axes of 
roll, yaw, and pitch. As regards surface targets on 
land or sea, it involves measurement of position, 
velocity, and acceleration in two degrees of freedom, 
specifically in the direction of range and of azimuth. 
For air targets the additional degree of freedom in 
the direction of altitude is added. 

For aerodynamic missiles, that is, missiles which 
derive their control from the forces exerted by the air 
mass on aerodynamic surfaces or on the missile’s 
structure, there must also be considered and meas- 


157 


158 


INSTRUMENTATION 


ured properties of the air mass; wind speed and direc¬ 
tion as a function of altitude are vital. Temperature 
and dew point may also be significant, especially if 
experiments should be abortive because of icing of 
control surfaces. With missiles of very high velocity, 
the relative position of these two properties may 
interchange, temperature and humidity becoming 
more important than wind velocity, since the speed 
of sound in air is inversely proportional to the abso¬ 
lute temperature. With missiles approaching sonic 
velocity it is probably more important to know the 
temperature of the medium and, therefore, the veloc¬ 
ity of sound and the Mach number of the missile than 
the ground speed of the air. With reaction-guided 
missiles, w T here transverse accelerations are imparted 
by the reaction of a laterally directed jet, it is possible 
that the whole problem of air measurement both as to 
wind velocity and temperature may become insig¬ 
nificant. 

8 2 2 Methods Used 

In the Division’s program all the attacks w^ere 
made against stationary targets. In general the posi¬ 
tion of the missile was determined from ground 
measurement. Observation from the ends of a base 
line can provide continuous information as to the 
altitude, range, and azimuth position of the missile 
during its flight. Such observations were made by 
phototheodolites, by conventional view cameras 
when the work w r as done at night so the shutters 
could be left open, and by specially constructed slit 
cameras. The angular position of the missile with 
respect to its three principal axes was in general 
recorded by means of missile-borne motion-picture 
cameras viewing the terrain approached. 

In general, the derivatives of the positions of the 
missile, both spatial and angular, w^ere not measured. 

In addition to the determination of the spatial 
position of the missile from the ground, supplemen¬ 
tary measurements were made with bomb-bay mo¬ 
tion-picture cameras in connection with the dirigible 
high-angle bomb project. These instruments pro¬ 
vided, perhaps, the most valuable means of obtaining 
data suitable for rapid analysis. If unaccelerated 
flight of the aircraft after release of the missile is 
assumed, they produce a ground projection of the 
trajectory from the point of view^ of the bombardier. 

Performance of the control system was recorded by 
means of motion-picture cameras, by chronographic 
recorders located in the missile, and by radiosonde. 


The data obtained from these instruments was in¬ 
valuable. The problem involved in film and record 
recovery and their subsequent analysis is not, how¬ 
ever, to be underestimated. A flight of one of the 
glide bombs of the Washington Project normally 
occupies approximately five minutes. With a motion- 
picture record taken at 16 frames per second, the 
analysis of 4,800 frames of motion-picture record is 
involved for every drop. Nor is 16 frames per second 
too fast a speed for recording such data. Indeed, 
during transient disturbances due to gusts, fading of 
the radar signal, or to other causes it is hardly ad¬ 
equate. Radiosonde techniques possess the possible 
advantage of producing the data in continuously 
plotted curves, a much more readily analyzed form 
than the moving picture film. Furthermore, the prob¬ 
lem of film recovery, which the Division never com¬ 
pletely solved for any of its projects, does not exist 
with this method. It requires, however, an additional 
radio lipk which must be carefully coordinated with 
other transmission in the vicinity if interference is to 
be avoided. 

8 2 3 Other Possible Methods 

The use of two phototheodolites to determine the 
spatial motion of the missile can give reasonably pre¬ 
cise data as to the trajectory of the flight. The addi¬ 
tion of a third phototheodolite gives multiple redun¬ 
dancy and increases the precision. Furthermore, it 
greatly increases the probability of a complete record 
throughout the flight from two instruments. The 
problem of coordinating the operation of an aircraft 
and tw^o or three phototheodolite stations on the 
ground is acute. Radio communication must be of 
very high quality—higher than that usually found in 
portable equipment. Time synchronization is essen¬ 
tial, and the development of a routine procedure 
similar to target tracking in coast artillery practice 
seems vital. Such a procedure is very difficult to fol¬ 
low where aircraft maneuvers are involved. In experi¬ 
mental work it is desirable that subordinate vari¬ 
ables, such as the crab angle of the dropping aircraft, 
be minimized. The direction of the bomb run is 
therefore not easily scheduled, especially if the winds 
aloft are variable. It is probable that radar tracking, 
using one or more such instruments as the SCR-584, 
should be invoked either in place of, or as an adjunct 
to, visual tracking by means of phototheodolites. 
This technique is particularly attractive since it 
offers a ready means of obtaining with some precision 




MISSILE-BORNE INSTRUMENTS 


159 


first and second derivatives of the spatial position of 
the missile. 

Angular motion of the missile is probably better 
recorded through gyroscopic instruments than from 
motion-picture records of the terrain. Three gyros 
can measure the angular position of the missile about 
each of the three axes. Quantitative take-off has not 
yet been fully developed, but it would appear that 
appropriately shaped condenser plates on the inner 
gimbal frame could be used in a circuit of reasonably 
high frequency so that the change in capacitance in a 
bridge circuit with changing relative position of the 
missile and gimbal frames could be determined. Such 
a take-off would impose no friction on the gyro. 

Rate gyros can measure directly the angular veloc¬ 
ity about each of the three axes. Such gyro instru¬ 
ments are already in an advanced state of develop¬ 
ment. 

Measurement of angular acceleration about the 
three axes poses a different problem. Angular ac¬ 
celerometers are required. An instrument comprising 
a polar moment of inertia, elastically mounted, would 
appear to be easily developed. Care must be taken 
to see that the natural period of such an instrument 
is well below that of any frequencies likely to be 
encountered. Strain gauge techniques for propor¬ 
tional takeoff are suggestive. 

The experience of the Division indicates that as 
the velocities of the missiles are increased, the meas¬ 
urement of velocity and acceleration, both linear and 
angular, will become crucially important. 

8 2,4 General Principles 

In addition to proving the vital necessity of ad¬ 
equate instrumentation, certain basic principles of 
instrumental techniques have appeared from the 
Division program. Derivatives, where required, (and 
as has already been indicated their importance is 
likely to increase) should be measured directly. The 
differentiation of experimental data is subject to 
error, whether it is done graphically or by a calculat¬ 
ing circuit. Physical measurements, particularly in 
the fields associated with guided missiles, are likely to 
yield fluctuating results. Their differentiation magni¬ 
fies these fluctuations, and the resulting derivative 
may be seriously false. Smoothing of the data can be 
employed, as in damping of instruments or by 
smoothing filters in the measuring circuit. Such cor¬ 
rective measures, however, are inherently processes 
of integration. A careful study of time constants and 


of frequency of variation is required if specious re¬ 
sults are to be avoided. 

Where different data are to be compared directly, 
it is desirable that they be recorded with the same 
instrumentation. In certain missiles (see Chapter 4) 
it is significant to measure the angle between the 
heading of the fuselage of the tnissile and the tangent 
to the line of flight. If the line of flight is determined 
from ground recoils, either by photo theodolites or 
otherwise, and the heading of the fuselage is deter¬ 
mined, for example, by a missile-borne motion- 
picture camera, the correlation of these records is 
tedious and can be ambiguous. The alternative would 
be a missile-borne instrument which would continu¬ 
ously measure and record the relative direction of the 
airstream with respect to the axis of roll of the mis¬ 
sile. This poses a problem in aerodynamic design. No 
very suitable angle-of-attack instrument has yet 
been developed. 

It would appear that radiosonde, particularly with 
a plotting board which would yield a graphical record 
of the missile-borne instrumentation is a better ap¬ 
proach to the problem than missile-borne motion- 
picture cameras. Time coordination would appear to 
be simpler, and the whole problem of film recovery is 
avoided. 

83 MISSILE-BORNE INSTRUMENTS 
8,3,1 Glide-Bomb Cameras 3 ' 4 

The initial camera installation on the glide bombs 
of the Washington Project consisted simply of stand¬ 
ard 16-mm motion-picture cameras with the spring- 
motor drive replaced by electric motors powered 
from the missile’s storage batteries. This camera was 
bore-sighted so that the optical axis would be tangent 
to the line of flight. Four miniature lights were so 
placed as to appear in the field of view of the camera 
but not so as to obstruct the portion of the terrain 
immediately adjacent to the principal point. These 
lamps were marked R, L, D, and G and were con¬ 
trolled through relays to light when the indication 
from the radar called for a right or left turn, dive, or 
an extension of glide. 

This somewhat qualitative record was soon im¬ 
proved by providing miniature instruments which 
indicated the output of the radar in milliamperes 
from the differential amplifier which biased the gyro 
coil (see Sections 1.6 and 1.7). An attempt was made 
to establish a standardized instrumentation scheme, 



160 


INSTRUMENTATION 


and the Division made a contract with Eastman 
Kodak Company to develop a recoverable motion- 
picture camera which would photograph the field of 


view along the line of flight together with the radar 
output meter and a mechanical indicator of the ele- 
von position. The very large number of variables 



Figure 1 . Camera installation in Pelican. (A) Internal camera, (B) external camera. 










NOSE CAMERAS FOR HIGH-ANGLE DIRIGIBLE BOMBS 


161 


capable of measurement, however, precluded the 
standardization of a single instrumentation scheme. 
Furthermore, GSAP (gunsight aiming point) cam¬ 
eras were more readily available from the Services 
than new cameras of the type developed under the 
contract. 

The method finally used on most of the flights 
placed two GSAP cameras within the fuselage of the 
missile (Figure 1). One camera photographed the 
field of view along the line of flight through a train of 
mirrors penetrating the fuselage skin. The other 
camera photographed an instrument panel contain¬ 
ing the elevon position indicators and instruments 
appropriate to measure the parameters of major in¬ 
terest. The film drive mechanisms of the cameras 
were mechanically connected to assure synchronism, 
and a cueing light in the field of the camera which 
viewed the terrain in series with a similar light of the 
instrument panel provided means of tying the film 
records together. Steel casing of 34-i n - plate lined 
with sponge rubber sufficed to protect the cameras 
from damage in crash landings. 

The data obtained from the internal camera (that 
recording the instrument board) was of predominant 
importance. The data from the external camera (that 
recording the terrain along the line of flight) was of 
qualitative value only. GSAP cameras are not pro¬ 
vided with fiducial marks which record sharply at the 
focal plane. The somewhat blurred outline of the 
frame has to be used. Since the data required from 
this instrument, however, comprised only the bearing 
of the target with respect to the line of flight, inaccu¬ 
racies remote from the principal point of the frame 
were not important. 

84 NOSE CAMERAS FOR HIGH-ANGLE 
DIRIGIBLE BOMBS 5 

The early high-angle dirigible bombs were made 
approximately the same size as a standard 1,000-lb 
GP bomb, but were fabricated of sheet metal and 
loaded with lead. The purpose of this construction 
was to provide room for instrumentation and later 
(see Chapter 2) to provide room for television. The 
instrumentation consisted of a standard 16-mm 
motion-picture camera. By means of partially sil¬ 
vered mirrors the image of the sweep second-hand of 
a stop watch was projected onto a ring immediately 
surrounding the central field of view. Small sectors at 
the top and bottom of the frame carried records of 
the position of the aerodynamic surfaces. 


This method of instrumentation was followed 
throughout the dirigible high-angle bomb program, 
including Felix. In the early work the cameras were 
recovered by parachutes ejected by squibs fired by 



Figure 2. Sample frames from Gulf nose camera. 





162 


INSTRUMENTATION 


barometric switches a few seconds before impact. 
Parachutes proved unreliable; in any event, the cru¬ 
cial information is obtained in the last few hundred 
feet of the flight. 

Camera cases were built of welded %-in. cold- 
rolled plate to protect the film against destruction 
at impact. 

The steel camera casing was not adequate to pro¬ 
tect the camera, nor indeed was it expected to be. 
The impact velocity of a dirigible high-angle bomb 
dropped from 15,000 ft is in excess of 800 ft per sec¬ 
ond. Even in moderately soft ground decelerations of 
several thousand g can be expected. Much of the film 



Figure 3. Nose camera assembly for high-angle 
dirigible bomb. 


was recovered (Figure 2) but almost never in one 
piece. The data were available, but too much credit 
can hardly be given to the contractor’s personnel for 
patience and perseverance in transcribing and ana¬ 
lyzing 16-mm motion-picture film in fragments. 

In the Roc program a similar problem confronted 
the investigators. For Roc-1 the cameras were 
mounted externally behind the dive brakes on the 
empennage. Their function was to photograph the 
field of view, consisting of a light target (see Chapter 
4). In addition, a telltale light between the wings in¬ 
dicated the operation of the rate gyro which limited 
the rate of roll of the missile. In this position, the 
cameras (Figure 3) were somewhat protected by the 
empennage structure of the missile; nevertheless, 
much film was lost. 

To correct this difficulty this contractor designed 
an armored magazine aimed to protect the film (Fig¬ 


ure 4). Initial tests with this type of camera showed 
that the film was protected within the cassette from 
tearing or breaking. The camera was dashed to 
pieces on impact, and the portion of the film in the 
shutter mechanism unwound and exposed much of 
the film contained in the cassette. A spring-loaded 
knife was mounted at the opening of the armored 
magazine. The shock of impact tripped the knife, 
shearing off the portion of the film which had been 
exposed and wound within the armored casing. The 



Figure 4. Cassette with shearing knife for Roc mis¬ 
sile-borne motion-picture camera. 


knife also helped to close the slot and prevent light 
from fogging the records. 

8 5 GROUND INSTRUMENTATION 
8,5,1 Phototheodolites 6 

Phototheodolites were used in connection with the 
high-angle dirigible bomb developmental work at 
Eglin Field. This use, however, was not successful 
because of the difficulty of coordinating the activities 
of the airship crew with the phototheodolite op¬ 
erators. 

In connection with the Roc project, however, they 
proved very valuable. The instruments are compli¬ 
cated and need to be carefully maintained. Well 
maintained, and in the hands of competent oper¬ 
ators, they can yield accurate data for missiles which 
are visible and whose speeds permit manual tracking. 
With two phototheodolites located at the end of a 
reasonably long base line, say 35,000 ft, the precision 
of location of any point on the trajectory is probably 
of the order of 50 ft. 


















GROUND INSTRUMENTATION 


163 


The use of three instruments giving three pairs of 
data would reduce the error by the square root of 3. 
A further advantage, however, would obtain. With 
two instruments, failure to track by one ruins the 
trajectory from that point onward. With three in¬ 
struments, the probability of a failure which would 
frustrate the reconstruction of the trajectory is re¬ 
duced to Care must be taken to insure that the 
bombing run and the flight path of the missile do not 
require the operator to look into the sun. Further- 


Figure 5. Washington cinema recording theodolite. 

more, a path of the missile over the zenith of any 
observer is almost impossible to track. 

The data obtained from the phototheodolite con¬ 
sists of 35-mm motion-picture film with the image on 
the missile photographed against a fiducial scale. 
Superposed on the same frame are the readings of the 
azimuth and elevation scale of the instrument. The 
record on the film thus consists of the bearing of the 
telescope in azimuth and in elevation with the error 
due to faulty tracking preserved. A time interval cir¬ 
cuit periodically photographs a serial number on the 
frame of all instruments, synchronizing them at that 
point. 



The data are thus presented in a form suitable for 
analytical, rather than for graphical, analysis. It ap¬ 
pears that a punched-card or punched-tape comput¬ 
ing machine technique would be valuable with the 
instrument. 

A simplified method of reducing the data was de¬ 
veloped by the contractor. 7 This method, easier to 
apply than the standard method of the technical 
manual, gives somewhat improved accuracy. 



Figure 6. Unattended K-24 camera station. 


8 5 2 Open Cameras 

Where experimental flights are made at night and 
the missile carries a flare, qualitative data which may 
have some quantitative significance can be obtained 
by ordinary view cameras located in the plane of the 
target and the bomb run and on its flank. Indeed if 
the cameras are provided with fiducial scales, good 
quantitative data may be obtained with them, using 
the technique well known to astronomers. 

Such quantitative refinement was not employed by 
the Division’s contractors. The use of open cameras, 
however, furnishes a rapid means of getting an over¬ 
all assessment of the trajectory. It is particularly 
valuable with homing missiles as it discloses to a first 
order of approximation whether the control is oscilla- 








164 


INSTRUMENTATION 


tory, damped, or unstable. Figures 12 and 13 of 
Chapter 4 present data of this type. 


8 5 3 Synchronized Ground Cameras 3 

Recording the motion of an object by simultaneous 
tracking from two base-end stations can be precise 
only if the instruments are carefully maintained. The 
optical axis must be carefully aligned with the azi¬ 
muth and elevation scales, and these must be free 
from backlash. Section 8.5.1 describes the use of 
instruments carefully designed to avoid these 
errors. 

In the early work of the Washington Project spe¬ 
cial theodolites were constructed, designed so that 
the azimuth and elevation scales could be photo¬ 
graphed from a single point of view. A stop watch 
was placed near the instrument scales and photo¬ 
graphed with them. At an oral signal the motor- 


driven cameras and stop watches were started, track¬ 
ing by both base-end operators having been estab¬ 
lished (Figure 5). 

Thi§ method assumes that tracking is perfect, 
since no photograph of the missile is obtained. Anal¬ 
ysis of the records, as in the phototheodolite method, 
leads to redundance in the apparent altitude of the 
missile. Without means of correcting for errors in 
tracking, discrepancies of several hundred feet in 
missile elevation were not uncommon. 

To eliminate these errors a photogrammetric 
method was evolved. Two AAF K-24 cameras were 
mounted at base-end stations. These cameras have 
motor-driven film transport mechanisms and an 
electrically tripped shutter. The minimum cycling- 
time is 0.3 second. The pictures are 5 in. by 5 in., and 
the magazine carries about 120 exposures. Lenses of 5- 
to 12-in. focal length were used. The shutters and 
transport controls were wired in series so that strict 
synchronism was assured. An intervalometer, tuning- 



Figure 7. Bomb-trajectory camera. 





GROUND INSTRUMENTATION 


165 


fork controlled, produced the impulses to trip the a photogrammetric process. The K-24 camera is pro¬ 
shutters. vided with fiducial marks which produce an accurate 

These cameras were left fixed during a trajectory coordinate system on the film. A transparent overlay 
determination (Figure 6). Thus the analysis became of 0.005-in. cellulose sheet was laid successively over 



Figure 8. Frontal (right) and profile (left) traces of typical Razon drop made with bomb-trajectory camera. 




166 


INSTRUMENTATION 


each exposure and the position of the missile pricked 
on it. With the aid of a watchmaker’s glass a preci¬ 
sion of approximately 0.003 in. could be obtained, 
which resulted in an overall precision of about 10 ft 
in the elevation of the missile. 

0 

8 5 4 Slit Cameras 8 

The phototheodolite technique described in Sec¬ 
tion 8.5 poses problems in coordination. In order to 
obtain trajectory information more rapidly a special 
trajectory camera was developed by Gulf (Figure 7). 
This instrument is panoramic in principle. The cam¬ 


era is fixed during a trajectory, and successive ex¬ 
posures are made on a single film. In order to avoid 
overexposure a moving mask successively uncovers 
those zones of the film where the image of the missile 
will appear. The film is held in a cylindrical frame 
concentric with the axis of tracking. A chronometer 
momentarily closes the shutter, giving a time record 
as well as the spatial geometry of the trajectory. In 
addition, the instrument having been developed for 
Azon and Razon, pilot lights controlled from a radio 
receiver, recorded the instructions given the missile. 
Figure 8 shows typical frontal and profile traces of a 
Razon drop. 



Chapter 9 

MISCELLANEOUS CONTROL SYSTEMS 


INTRODUCTION 

I n addition to the four integrated systems of 
guided missiles discussed in Part I of this report, 
the Division undertook the development of a small 
number of separate control devices. All but two of 
these control devices were calculated to produce 
homing flight of the missile for test purposes only. 
There were two reasons for adopting this course of 
action: (1) homing devices suitable for combat appli¬ 
cation were in general complicated and difficult of 
procurement; (2) as an interim measure during the 
development of the missile, it was essential that a 
homing device so simple in concept that its reliability 
was certain could be available to test fundamental 
homing qualities of the missile and other elements 
of the servo system. 

There was also the possibility that a control device 
might appear in a laboratory not connected with the 
development of missiles. It was conceivable that a 
profitable development could be undertaken under 
these circumstances. The Division undertook two 
programs in this category: the organic homing system 
and the to-and-fro scanner. Neither of these devices 
reached a successful degree of development, although 
there is some evidence that had it been possible to 
achieve closer integration between the group engaged 
in the development of control devices and those 
working on the other portion of the servo system, 
each of them might have been carried at least to suc¬ 
cessful test, if not to actual combat application. 

Until a system of analysis appears which will indi¬ 
cate clearly and reliably the interrelationship be¬ 
tween all the elements which comprise such a guided- 
missile system, separate development of components 
followed by their assembly can only be regarded as 
most hazardous. 

92 HIGH-ANGLE PHOTOELECTRIC 
HOMING BOMB 1 

9,2,1 General 

As has been discussed in Chapter 3, the Felix bomb 
under development by the Massachusetts Institute 
of Technology was conceived as a high-angle bomb 


having the minimum departure from the standard 
geometry of the 1,000-lb GP bomb. Chapter 2 dis¬ 
cussed the cooperative investigation made by MIT 
and the Gulf Research and Development Company 
in the evolution of a dirigible high-angle bomb. This 
cooperation was continued in the development of 
Felix. Gulf concentrated on the development of a 
high-angle bomb demonstrably capable of homing, 
MIT on the sensory or target-seeking device. For 
preliminary work it was impossible and indeed un¬ 
desirable to use thermosensitive elements to detect 
the target. For the experimental purposes of proving 
the capability of the missile to home satisfactorily, a 
photoelectric target seeker seemed much more desir¬ 
able. Photoelectric cells of high sensitivity were plen¬ 
tifully available. By performing tests at night, a very 
high degree of contrast between the target and the 
background surrounding it could be assured. 

The experimental bomb was constructed of sheet 
metal, like the early Azon and Razon bombs, and 
was provided with a cylindrical shroud substantially 
at the center of gravity to increase its maneuverabil¬ 
ity (Figure 1). The empennage was made octagonal 
as a result of the work done in the wind tunnel at 
MIT in connection with roll torques and cruciform 
structures with simultaneous yaw and pitch. (See 
Chapter 2.) 

The missile was stabilized in roll in the same man¬ 
ner as the Azon and Razon. Absolute roll stability 
in a homing missile of symmetrical structure is prob¬ 
ably unnecessary. In fact, it will be shown in Sections 
9.3 and 9.4 that Roc was made to home successfully 
without being so stabilized. All that is necessary is to 
limit the rate of roll in such a manner that the sen¬ 
sory device which detects the target can cause ap¬ 
propriate operation of the control surfaces to produce 
a lift in the direction to correct the error. This means 
that the scanning period, with the accompanying lag 
in computing circuits and operation of servo link, 
must be short enough to permit only a negligible 
change in the attitude of the missile about its roll 
axis. In this project, however, the determination of a 
suitable limit for rate of roll seemed undesirable since 
absolute control of roll had already been established 
by the contractors in connection with the Azon and 
Razon projects. 


167 


168 


MISCELLANEOUS CONTROL SYSTEMS 


Missiles with fixed wings, such as the lift shroud 
provided for the photoelectric target-seeking bomb, 
or with bomb casing itself acting as a wing, fly with an 
appreciable and varying angle of attack. A scanning 
device rigidly mounted in such a structure will have 
a field of view centered around some rigid axis in the 
missile, which in general is not along the line of flight; 
that is, a homing device simply will not look where 
it is going. For this reason, a portion of the optical 
system of the photoelectric scanner was mounted on 
a wind-vane-actuated support, so that the axis of 
scan was always parallel to the axis of flight. Poten- 



Figure 1. Photoelectric target-seeking bomb. 


tiometers mounted on the control surfaces provided 
a feedback to cutoff so that the position of the rudder 
and elevators satisfied the equation 

8 = - kiO (1) 

where 8 is the rudder or elevator displacement and 0 
is the course error in the corresponding sense. 

9 2 2 Dynamics of Flight 

Figure 2 represents a portion of the trajectory of a 
homing missile. At the instant under consideration 
the missile has a velocity along the vector V to pro¬ 
duce an expected miss h. The homing device, which 
is assumed to look along the velocity vector, meas¬ 
ures the course error 0. Under the system of control 
selected for the photoelectric target-seeking bomb, 
the servomechanism causes the rudders to deflect 
proportionally to 0. 

For small values of 0, such that tan 0 is negligibly 
different from 0, the lift L is proportional to the angle 
of yaw 4'/, which in turn is proportional to the rudder 
displacement 8. 


L = at (2) 

t = c 2 8 (3) 

8 = h0 (4) 

L = C\C 2 ki0 

= -k0 (5) 

Gravity being neglected, the lift is balanced by two 
accelerations, centripetal and yawing, so that 

— k0 = k z 0 + kit (6) 


Under the assumption of linearity between t and 0 
this reduces to 

k$0 kz0 -|- k0 — 0 

0 = fa k,t,U sin (|/|r - + P (7) 

the well-known damped oscillation. 

The validity of equation (6) is dependent on the 
assumption of linearity expressed in equations (1) 
to (5) and also upon the assumption of secularity of 
such parameters as the range R and the velocity. 
Furthermore, there must be no time lags in the sys¬ 
tem of control. Actually such time lags always exist 
so that the real control equation is 

8 = - k0f{t) (8) 

The presence of lag introduces negative damping. 
As the missile comes on course, the rudders should be 
in neutral under the control regime selected. If they 
lag, however, then an aerodynamic moment appears 
which causes an overshoot. Energy is fed into the 
system in an amount equal to the summation of the 



Figure 2. Projection of homing trajectory. 


overshoot multiplied by the moment producing it. 
Only one thing appears tobalancethissource of energy, 
namely the aerodynamic damping of the missile. 
Transient wind-tunnel studies are needed for its de¬ 
termination. 

Photoelectric Homing Device 

Scanning System 

The preceding section showed the necessity of pro¬ 
portional control. The optical system, including the 






HIGH-ANGLE PHOTOELECTRIC HOMING BOMB 


169 


method of scan, was designed to attain it. An objec¬ 
tive lens focused the target image on a scanning disk. 
The scanning disk was provided with slits mutually 
at right angles which periodically swept the field of 



Figure 3. Optical system of scanning unit. 


view in an approximately horizontal and vertical 
direction. The target, therefore, appeared as a single 
flash of light through the slit which ultimately deter¬ 
mined its vertical or its horizontal position within the 
field of view. A triple condensing lens focused the 
spot of light on the cathode of the photocell irrespec¬ 
tive of the position of the spot within the frame. 


Figure 3 shows the limit of displacement of the 
objective lens in one direction due to the operation 
of the wind vane. The wind vane controlled a racking 
system similar to the rising and falling front of a 
view camera. Racking in the pitch direction was ac¬ 
complished by two wind vanes connected to the rack 
mechanism through a differential which averaged the 
direction of the wind stream on each side of the bomb. 
Racking in the yaw direction was accomplished by a 
single wind vane well below the body of the bomb to 
avoid turbulence. The differential construction was 
impossible in the yaw direction because the presence 
of a wind vane on top of the bomb would have inter¬ 
fered with hanging the bomb from the usual shackle. 

The slits of the scanning disk swept the image area 
alternately up and down and then left and right. The 
scanned image of a point source produced a pulse 
output of the phototube. This pulse was amplified 
and fed into the grid of a thyratron (Figure 4). The 
plate circuit of the thyratron was closed by a com¬ 
mutator at the beginning of the scanning cycle, thus 
setting the tube up to conduct at the instant a pulse 
was received. The tube remained non-conducting un¬ 
til the pulse on the grid triggered it, at which time it 
remained conducting until the commutator broke the 
plate circuit at the end of the scanning cycle. 

A voltage thus appeared across the plate resistor 
of the thyratron as a square wave, the length of 
which measured the position of the point of light in 
the,field of view. This square wave was filtered and 
amplified to operate relays controlling the rudder and 
elevator servomotors. The final stage of the control 
amplifier was biased by a voltage fed back from po¬ 
tentiometers driven by the servomotors operating the 
control surfaces. Thus, any given length of pulse 
would establish a voltage level in the control ampli¬ 
fier. This voltage would be extinguished by the bias 
voltage from the control surfaces, and proportional¬ 
ity was obtained. 


FLARE 
ON GROUND 



Figure 4. Schematic diagram of scanning system 


































170 


MISCELLANEOUS CONTROL SYSTEMS 


Phototube and Amplifier 

The phototube selected was a CE-l-AA with a 
sensitivity of approximately 300 microamperes per 
lumen. The pulse output of this tube was amplified 
through two pentode and three triode stages. The 
pulse amplifier was resistance coupled, with time 
constants of the coupling circuits selected to give 
negligible phase distortion in terms of the scanning 
frequency (approximately 40 c). A single photocell 
scanned both up-down and right-left directions, the 
output of the pulse amplifier firing a type 2050 thyra- 
tron. Alternate pulses from the thyratron were 
switched by the commutator to the left-right and up- 
down channel so that the actual scanning frequency 
is approximately 20 c. 

With the target directly on course, the pulse will 
appear in the center of the scanned field, and the 
rudders and elevators will assume the position re¬ 
quired to maintain that course. The correct heading 
depends upon the continuous reception of a light 
signal from the target. Failure of the signal due to the 
extinguishing of the target flare or the failure of the 
phototube would cause the rudders and elevators to 
go to their extreme position, say left and down. 

In the first few seconds of flight until the bomb has 
nosed over on its parabolic trajectory, the target is 
not in the field of view. An auxiliary relay was pro¬ 
vided to hold the rudders and elevators in the neutral 
position until the target first appeared in the field of 
view. It then closed and sealed itself in, making the 
servomotors operative. This system can hardly be 
considered “fail safe.’ 7 The relay which centers the 
rudder and elevator should not seal in but should be 
allowed to return to the centering position after a 
reasonable time delay following the loss of a target 
signal. 

For the purposes for which this homing device was 
intended, this was not an important feature. It 
should, however, be incorporated in any combat ap¬ 
plication of a homing device. Loss of signal should 
not result in gross errors, that is, errors greater than 
would have been attained had there been no attempt 
at automatic homing. 

Radiosonde 

Azon and Razon were checked during the early ex¬ 
perimental phases of their development by means of 
motion-picture cameras located in the nose, photo¬ 
graphing the terrain toward which the bomb was 


falling. With the photoelectric target-seeking bomb, 
all operations were planned to take place at night, 
and such photographic techniques were imprac¬ 
ticable. Accordingly, a simple radio transmitter was 
installed in the bomb to transmit to a ground station 
the output pulse, the length of which measured the 
position of the target with respect to the line of flight. 
The received pulse was applied to the vertical deflec¬ 
tion circuit of an oscilloscope through an electronic 
switching circuit which produced reversal of alternate 
square-wave pulses. Thus the length of positive pulses 
on the oscilloscope screen measured the bearing of 
the target in one coordinate (e.g., azimuth), and the 
length of the negative pulse measured the bearing in 
the orthogonal direction. These measurements were 
recorded by means of 16-mm moving-pictures (Fig¬ 
ure 5). 

9 2 5 Dropping Test 

Five missiles were dropped at Eglin Field in De¬ 
cember 1943. Of these five missiles, one failed; the 
remainder made scores of 116 ft, 89 ft, 94 ft, 243 ft, 
and 47 ft. 

The target consisted of a 1,000,000-candlepower 
pyrotechnic flare. At an altitude of 4,000 ft, the re¬ 
sulting intensity of the photoelectric target-seeker 
was approximately 34 times the threshold of opera¬ 
tion. In addition to the radiosonde system just men¬ 
tioned, records were taken with bomb-bay cameras 
giving the ground projection of the bomb from the 
point of view of the aircraft. Slit cameras approxi¬ 
mately 6 miles on the flank of the trajectory recorded 
its profile (Figure 6). 

Frame-by-frame analysis of the records from the 
radiosonde apparatus gave a continuous plot of the 
apparent bearing of the target along the line of flight 
(Figure 7). A similar analysis of the bomb-bay cam¬ 
era gave the approximate ground projection of the 
trajectory (Figure 8). 

Section 9.2.3 showed that a control system, in or¬ 
der to produce nonoscillating flight, must develop 
lift forces which are proportional to the time deriva¬ 
tive of error in heading, as well as to the error itself. 
The records, just given for bomb No. 83, which are 
typical of the four successful missiles, failed to show 
any signs of such oscillation. The yawing and pitch¬ 
ing of such a missile as this to produce lift forces in 
the azimuth and range direction must be accom¬ 
panied with sufficient aerodynamic damping to absorb 
enough energy to make the system nonoscillatory. 



HIGH-ANGLE PHOTOELECTRIC HOMING BOMB 


171 



Figure 5 


Successive frames from radiosonde oscilloscope record 



























172 


MISCELLANEOUS CONTROL SYSTEMS 



seeking. The extreme scarcity of radar target-seeking 
equipment made it wholly undesirable to use this 
means of intelligence during the developmental stages 
of the missile itself. As a portion of their contract for 
the development of the missile, the Douglas Aircraft 
Company undertook the development of a photo¬ 
electric target-seeking device to prove the homing- 
property of the missile. 

The scanning system consists of a cloverleaf photo¬ 
cell having four cathodes, each occupying approxi¬ 
mately one quadrant sector of a disk 13 ^ in. in diam¬ 
eter (Figure 9). This photocell is located in the focal 
plane of a scanning lens which is mounted eccentri¬ 
cally with respect to the four cathodes. The eccentric 
lens with its housing is rotated at 1,800 rpm concen¬ 
trically with the photocell axis. 

With this scanning system a target flare describes 
a circle as the objective lens rotates around the cen¬ 
ter of the quadrant-cathode array. If the target is 
dead ahead, the circle will lie for an equal time inter¬ 
val-in each of the quadrants. If the target is off course, 
the duration of its excursion within one of the quad¬ 
rant sectors will be greater than its excursion on the 
diagonally opposite one. The time difference between 
the length of excursion in diametrically opposite 
cathodes is a measure of the error in heading. 

A two-channel preamplifier accepts the output of 
two diagonally opposite cathodes, amplifies them to 
a suitable level, and in a saturated stage clips them. 
The result is a pair of pulses of equal amplitude but 
differing in length. These pulses are then subtracted 
and the output filtered and integrated. For each of 









QUADRANT PHOTOCELL TARGET SEEKER 


173 


the two channels, right-left or up-down, there then 
appears a d-c voltage which is a function of error in 
heading. For small errors in heading, the relationship 
is nearly linear, so that the output of the photocell 
amplifier for small errors in heading can be consid¬ 
ered to be: e = Kd, where e is the output of the am- 



Figure 8. Projected path of Gulf photoelectric target¬ 
seeking bomb No. 83. 


plifier, K is a factor of proportionality, and 6 is the 
angular error in heading. This voltage is then fed to 
a differentiating-mixing circuit which develops a volt¬ 
age for driving the flap motor at a speed proportional 
to the course error and to its first time derivative. 
(See Chapter 4.) 

The missile is a four-winged device with individual 
servomotors driving full-span flaps on each of the 
wings. Consequently, each of the two channels has 
two outputs, and each of the opposite wings is con¬ 
trolled from one of the amplifier channels. The up- 
down channel (Figure 10) is biased by a feedback 
voltage from potentiometers mounted on the wing 


flaps; if the displacement of one flap advances ahead 
of that of its mate, a voltage is injected into the con¬ 
trol amplifier to extinguish this differential, thus 
maintaining control surfaces of this pair of wings in 
phase. 

No attempt is made to keep the control surfaces on 
the other pair of wings in step. In fact, a rate gyro¬ 
scope is provided on this channel to bias the speed 
of the wing motors to correct excessive rate of rota¬ 
tion. To repeat, no attempt is made to maintain an 
absolute roll of stabilization. Designation of the chan¬ 
nel, therefore, as “up-down” or “right-left” is for 
convenience only and has no physical significance. 
As the missile rolls during its flight, the different 
pairs of wings successively perform the function of 
providing lift in the range and in the azimuth direc¬ 
tions. It was well understood that such a target seeker 
was somewhat qualitative in its response. Figure 11 
shows that the output of the photocell preamplifier 
depends not only on the error in heading but also on 
the bearing. Further, it shows that the right and 



Figure 9. Section of quadrant photocell target seeker. 


down, or better, the x and y components, or the error 
in heading were not pure orthogonal components; 
that is, the voltage resulting from an error in heading 
of 5 degrees in azimuth varies with varying errors in 
heading in the range sense. These vagaries, however, 
were considered as not seriously affecting the signifi- 















































174 


MISCELLANEOUS CONTROL SYSTEMS 


cance of the experiments in proving the homing char¬ 
acteristics of the missile. 

Other difficulties with the system were, however, 
serious. The initial quadrant photocell was a CE-47 
which the Army Air Forces had made available for 
this project. Serious variations appeared in the sen¬ 
sitivity of the various quadrants, even within a single 
tube. In many cases these variations were beyond the 
capability of the preamplifier to limit and clip. Fur- 


Y SCANNER LENS ANO MOUNT 


CATHODES 


EZZL 



RIGHT - LEFT 


UP-DOWN 

CHANNEL 


CHANNEL 


RIGHT 

-LEFT 

RIGHT 

LEFT 


RIGHT 

QQ left | 
PHOTOCELL g] 

nn 


OUTPUT 


DOWN 


PREAMPLIFIERS 


w 

UP 


UP-l 

DOWN 

UP 

DOWN 


OUTPUT AMPLITUDES LIMITED 
AND CHANNEL PULSES 
INTEGRATED PER CYCLE 

MOTOR DRIVE AMPLIFIERS 


FOLLOW-UP 

CONTROL 


OUTPUT 

RIGHT AND LEFT CHANNELS 

e = K,(t,-t 3 )+ K 

GYRO ACTION 

UP AND DOWN CHANNELS 

• ■K,<Vt,)+K 4 $<Vt,>± 

FOLLOW-UP BIAS 


Figure 10. Block diagram of quadrant photocell 
target seeker and control circuit. 


thermore, the glass work of the envelope was seriously 
defective, the end of the envelope being so irregular 
that the path of the rotating circle of the target image 
was seriously distorted. This tube was finally replaced 
by a much better quadrant tube especially developed 
by the Farnsworth Radio and Television Corporation 
through the cooperation of Section 16.4, NDRC. The 
four cathodes of this tube were mounted in an en¬ 
velope having an end of optical glass optically pol¬ 
ished. Greatly increased sensitivity was obtained by 
photomultiplication. Each cathode quadrant had its 
individual assembly of multiplier cups and collector 


anodes so that the whole tube comprised four six- 
stage photomultipliers. 

The basic scheme of operation of the preamplifier 
involves the measurement of the length of the volt¬ 
age pulse arising from the time that the target image 
is on a particular quadrant. This requires that the 
sides of the voltage pulse shall not deviate materially 
from the perpendicular. The physical circuit cannot 
tolerate even finite discontinuities. It is impossible 
for a voltage to rise instantly across a condenser 
without the expenditure of energy at an infinite rate. 
An engineering compromise is effected by designing 
circuits which would tolerate reasonably rapid volt¬ 
age rises, and by incorporating into the power supply 
to the dynodes current-limiting resistors to provide 
within the photomultiplier itself a measure of AGC, 
so that the clipping action required from the pre¬ 
amplifier would not be excessive. Thus spurious 
lengthening of the pulse due to clipping farther and 
farther down on a trapezoidal wave was eliminated. 

•Test results with this homing device have been dis¬ 
cussed at some length in Chapter 4. In summary it 
can be said that the quadrant photocell target seeker 
gave good qualitative proof of the homing character¬ 
istics of Roc and a measure of quantitative support 
for the particular control regime selected. 


94 WIDE-ANGLE 

PHOTOELECTRIC SCANNER 3 

9-4,1 General 

The photoelectric target seeker just described had 
certain inherent limitations. The corrective signal in 
azimuth depended not only on the azimuth error but 
also on the error in range. Furthermore, as has been 
already mentioned (see Section 9.3), difficulties in 
procuring suitable quadrant photocells made the out¬ 
come of the project rather doubtful. To overcome 
both of these difficulties, the Division made a con¬ 
tract, OEMsr-1182, with the Fairchild Camera and 
Instrument Corporation to develop a wide-angle 
photoelectric scanner which would develop a signal 
suitable to control Roc through the servo link de¬ 
signed for it. 

Specifically, this requirement is for an output volt¬ 
age from the scanner such that when 

0 < 6 < 10 degrees 
e = kid 










































WIDE-ANGLE PHOTOELECTRIC SCANNER 


175 



Figure 11. Output of quadrant photocell preamplifier. Parameter <f> is measure of direction of total error. 

_ ^ _ L degrees range error 
^ an degrees azimuth error 


and when 

10 degrees < 6 < 40 degrees 
6 = E 

The general approach was to scan the landscape so 
that a luminescent target such as a pyrotechnic flare 
would produce a spot of light on the cathode of a 
single photomultiplier tube. The optical system was 
arranged so that the length of the pulse produced by 
the image of the target would be proportional to the 
error in heading for small angles and constant for 
larger errors. A cardioid aperture in the focal plane 
of a simple lens, if rotated at a constant speed, pro¬ 
duced a pulse of light the length of which was pro¬ 
portional to the error in heading of the missile with 
respect to a luminescent target. Such a system, how¬ 


ever, required good optics; the resolution had to be 
such that the size of the image of the target was small 
in comparison with the width of the cardioid slit at 
the smallest error angle which it was desired to meas¬ 
ure. Without resorting to elaborately corrected lepses, 
it was found impossible to attain so small a circle of 
confusion. For errors of the order of 1 to 2 degrees 
off course, the diameter of the target image was 
greater than the width of the cardioid aperture. 

The general system was reversed, therefore, and an 
aperture similar in shape to half a keyhole was made, 
such that the light from the target excited the photo¬ 
multiplier cathode for a period varying proportion¬ 
ately from of a scanning cycle with zero error in 
heading to V 24 of a scanning cycle for a 10-degree 
error. For errors in excess of 10 degrees, the excita- 






176 


MISCELLANEOUS CONTROL SYSTEMS 


tion period was constant at V 24 of the scanning 
cycle. 

The output of the scanning cycle was then made 
proportional to the dark time of the phototube. A 
commutator driven synchronously with the scanner 
analyzed the output voltage into orthogonal com¬ 
ponents and distributed them to the up, down, right, 
and left channels of the motor-control amplifier. The 
Roc missile is essentially a Cartesian structure. Cor¬ 
rective instruction to the control apparatus has to 
be given in terms of right, left, up, or down, or bet¬ 
ter, since the missile is not absolutely stabilized in 
roll, in terms of fxor ±y. The commutator, there¬ 
fore, must not only distribute the output of the scan¬ 
ner amplifier to the appropriate quadrant correspond¬ 
ing to the wing system on the missile, but must also 
resolve the output in accordance with the sine and 
cosine of the instantaneous roll attitude of the missile. 



Figure 12. Design of Fairchild aperture with ideal 

optics. 

9 4 2 Optics 

A synchronously driven shutter interrupts a light 
for a period proportional to the bearing off axis if it 
has a semicircular aperture which rotates about one 
end of its boundary diameter (Figure 12). The diam¬ 
eter of the semicircle in degrees is equal to the portion 
of the field of view in which proportionality is de¬ 
sired. A sector from the outer edge of the semicircle 
to the edge of the field of view gives constant dark 
time for errors in heading greater than the semicircle 
diameter. The scanning aperture thus has the shape 
of half a keyhole. For the signal to be of the character 
desired, the aperture (Figure 13) must be located in 
the focal plane of the lens and must rotate about the 
optical axis. Then, if the target is dead ahead, the 
image of the target will lie at the principal focus of 
the lens, which is the center of rotation, and no pulse 
will be transmitted through the aperture to the pho¬ 
tocell. Such a system, however, has a limited field of 


view; 35 degrees is approximately the limit of total 
angle attainable with a system of uncorrected lenses 
without experiencing serious aberration, with con¬ 
sequent increase in the size of the circle of confusion. 

In order to increase the angle of the field of view, 
a wedge prism was mounted in front of the objective 



lens. The angle of the prism was such as to bend the 
optical axis of the system so that the boundary gen¬ 
eratrix of the cone of scan was parallel to the axis of 
the objective lens. The whole optical system, then, 
prism, objective lens, and iris diaphragm, was rotated 
about the optical axis of the objective lens. This gave 
the optical system a squint so that at one instant it 
viewed the terrain from dead ahead to 40 degrees to 
one side. One-half scanning cycle later it viewed the 
terrain from dead ahead to 40 degrees in the reverse 
sense. A second effect of the prism was to move the 
principal focus of the entire optical system to a point 



Figure 14. Actual shape of keyhole scanning aperture. 

on the edge of the field of view in the focal plane at 
the end of a radius perpendicular to the dihedral of 
the prism. As the optical system was rotated by the 
scanning motor, a point dead ahead would be con¬ 
tinuously focused on the outer end of this radius. All 
other points in the field of view would describe con- 




































WIDE-ANGLE PHOTOELECTRIC SCANNER 


177 


centric circles around this point. The radius of these 
concentric circles was proportional to the error in 
heading. The total effect of the prism, then, was to 
increase the field of view and, by shifting the prin¬ 
cipal focus of the optical system from a fixed point 
at the center of the focal plane to a synchronously 
revolving point on its boundary, to require that the 
aperture (Figure 14) be reconstructed so as to cover 
substantially the entire diameter of the focal plane. 
Some distortion resulted from the use of the prism, 
and the aperture no longer consisted of a semicircle 
terminating in a wedge sector. 

The technique of producing the aperture is worthy 
of note. A shape simulating the ideal aperture (see 
Figure 12) to scale was cut out of black cardboard 
and mounted on a white screen. This shape was then 
floodlighted and photographed on a sheet of Kodalith 
film mounted in the focal plane of the scanner with 
which the aperture was to be used. This technique 
automatically compensated for the aberrations in the 
optical system. After the film was developed, retouch 
artists made the negative opaque, to give sharp 
boundaries to the aperture. The iris diaphragm is 
located at a focus conjugate with the cathode of the 
photomultiplier. 


9 4 3 Preamplifier 

The phototube was a type 931-A ten-stage photo¬ 
multiplier. With 1,000 v between the cathode and the 
ninth dynode, threshold sensitivity was approxi¬ 
mately microlumen, 0.001-footcandle intensity at 
the scanner. This threshold sensitivity is equivalent 
to a signal-to-noise ratio of about 2. 



Figure 15. AGC arrangement in phototube connec¬ 
tions. 



Figure 16. AGC characteristics of wide-angle scanner preamplifier. Response approximately proportional at 8 foot- 
candles. 















































































178 


MISCELLANEOUS CONTROL SYSTEMS 



Figure 17. Block diagram of wide-angle scanner circuit. 


The illumination intensity undergoes enormous 
gain during a missile flight. In a drop from 10,000 ft 
(14,000-ft slant range) it rises from 0.005 foot- 
candle at release to 4 footcandles at the beginning of 
the last second before impact, some 58 db rise in 
power level. It was therefore necessary that the pre¬ 
amplifier supplied by the phototube have as much 
AGC as was practicable (Figure 15). In addition it 
was necessary to limit the current from the dynodes 
at stages 5 and 7. The overall effect of the AGC am¬ 
plifier and the photomultiplier power-supply net¬ 
work was to give substantially flat voltage from 0.001 
footcandle to 4.0 footcandles. 

After they were amplified, the pulses were clipped 
to a fixed amplitude and then integrated to produce 
a triangular pulse, the amplitude of which was pro¬ 
portional to the excitation time of the phototube 
(Figure 17). The output of the integrator was fed to 


a differentiating amplifier, which produced a very 
high pulse at the trailing edge of the triangular wave. 
This pulse was used to trigger a square-wave genera¬ 
tor. The amplitude of the triangular pulse, passed 
through a rectifier, was applied to the square-wave 
modulator. 

The square wave produced by the generator was 
essentially a gate exactly 34 of the scanning cycle 
wide. The output of the square-wave modulator, 
therefore, was a pulse 90 degrees in width and having 
an amplitude proportional to the error in heading for 
small errors and constant for larger errors. This 
wave was suitably amplified and commutated. The 
commutator was phased mechanically with the scan¬ 
ner and the wing structure of the missile. A ladder 
filter in the output of the commutator provided 
smoothing so that the final output was substantially 
a d-c voltage. 



























































WIDE-ANGLE PHOTOELECTRIC SCANNER 


179 


UJ 




Figure 18. Commutation system and ladder filter for 
wide-angle scanner. 


In order to change the polar coordinate character¬ 
istics of information developed by the scanner to the 
Cartesian information required by the wing structure 
of the missile, the first element of the ladder filter 
(Figure 18) consists of a series resistor and a con¬ 
denser, the constants of which are so selected that 
the charging rate of the network closely approximates 
a sine function. Without these stages, the common 
0.1-megohm resistor and the 0.04-juf condenser, the 
x and y components of scanner output would be 


proportional to the instantaneous phase in roll of the 
missile rather than to the sine and cosine of this 
phase. Figure 19 shows the degree of approximation 
to the appropriate sine and cosine function obtained 
by the exponential function expressing the charge of 
the initial condenser in the output filter of the scan¬ 
ner system. Figure 20 shows the output of the pre¬ 
amplifier and wide-angle scanner combined. 



Figure 20. Output of preamplifier and wide-angle 
scanner. 



Figure 19. Approximation of sine function by con¬ 
denser-resistor combination. 


9 4 4 Test Results 

Six scanning devices were constructed in accord¬ 
ance with the principle above described. Three were 
tested with Roc missiles during the early summer of 
1944; the remainder, unexpended, have been turned 
over to the Air Technical Service Command. 

Considerable difficulty was experienced, due to er¬ 
ratic operation of the control surfaces when operated 
by the output of this device. As has already been 
stated, the control system for Roc (see Chapter 4) 






























180 


MISCELLANEOUS CONTROL SYSTEMS 


was built about a regime of control such that the rate 
of change of wing incidence angle was proportional 
to the error in heading and to its first time derivative. 
The error in heading was measured by the output of 
the amplifier associated with the wide-angle scanner. 
This output was then passed to a differentiating- 
mixing circuit for each of the two channels, right- 
left, or up-down, or preferably, + x — x, y — y. 
The output of the mixing circuit produced a voltage 
proportional to the input voltage and to its first time 
derivative. The factors of proportionality were such 
that the derivative component in the output was 
about five times the component due to input voltage. 
Thus the whole control system following the pre¬ 
amplifier was extremely sensitive to irregularities and 
fluctuations in the output voltage of the preampli¬ 
fier. 

Fluctuations can result from several causes. Those 
due to variation in light intensity can be eliminated 
by an appropriate AGC system; this was done. Those 
due to noise originating within the preamplifier cir¬ 
cuit and to interference from other radiating circuits 
in the missile are overcome with more difficulty. Cir¬ 
cuits with high grid impedances are much more sub¬ 
ject to both of these types of fluctuation than are 
circuits with moderate grid impedances. The pre¬ 
amplifier associated with the wide-angle scanner con¬ 
sisted of a triode stage, a pentode stage, and two 
triode stages. The grid impedances respectively were 
10 megohms, 5 megohms, 10 megohms, and 5 meg¬ 
ohms. Even beyond this point, grid resistors of higher 
value than 3 megohms were the rule. 

After the input circuit to the preamplifier is care¬ 
fully shielded and suppression filters are applied to 
the noise-generating equipment on the missile—gyro 
motors, wing-flap motors, and scanning motor—the 
output of the preamplifier associated with the wide- 
angle scanner for various errors in heading up to 10 
degrees is as shown in A of Figure 21. B shows the 
results of differentiating the output for 1 degree, 5 
degrees, and 10 degrees. The corresponding armature 
voltage is shown in C, and the flap displacement vs 
time resulting from the overall control system is 
shown in D. Three missiles were dropped with this 
control system. One of the three failed to operate 
for causes which could not be determined. The two 
remaining missiles scored errors of 142 ft and 372 
ft, correcting aiming errors of 2,400 ft and 900 ft 
respectively. 

Three drops constitute a wholly inadequate sample 
on which to base mature conclusions concerning a 


device of this nature. The project was relinquished 
in favor of other developments in connection with 
Roc. So far as photoelectric target seeking itself is 
concerned, this development seems an adequate solu¬ 
tion to the problem of providing quantitative re¬ 
sponse over a wide field of view. With some further 
engineering, directed, perhaps, toward simplification 
of the preamplifier, there appears to be no reason why 
it would not furnish adequate means of testing hom¬ 
ing missiles. 

9 5 ROLL STABILIZATION 

WITHOUT A FREE GYRO 4 

In Section 2.4 the problems of roll stabilization 
were discussed in detail. Early in 1943 the Division 
was seriously concerned over the threat to the Razon 
program due to failure of the roll-stabilization system 
then in use, and also wished to make a stabilizer 
available for Roc should it be needed. Electro- 
Mechanical Research, Inc., who had worked upon 
the problem, was asked to perfect its device. Eventu¬ 
ally Gulf and Douglas solved the problem for their 
specific missiles, but in the meantime some work had 
been done on the alternative device. 

Lacking a free gyro, this system reduced roll to a 
minimum value rather than eliminated it. This was 
in accordance with the thought (Section 2.4.1) that 
in cylindrical coordinates or for a homing system 
(Chapter 4) a slow roll is inconsequential. 

Theory of Operation 

The EMR device used conventional ailerons for 
control, the only unique feature being the method b} r 
which hunting was minimized. By means of a con¬ 
denser memory circuit, the position of the ailerons 
was reversed whenever the angular velocity was re¬ 
duced to an arbitrary fraction, for example, }/% of its 
original value. In the absence of time lags in the 
system, rolling should be reduced to zero in less than 
three cycles. In an actual mechanism, sustained oscil¬ 
lation will occur, the amplitude of which varies with 
the cube of the total time delay, with the size of the 
flaps, and inversely with the run-out time of the flaps. 

Ideal Case —No Time Delay 

Case 1. No initial acceleration. In this case control 
starts when the bomb has an angular velocity <j> but 
no initial accelerating torque. At t = 0 a motor starts 



ROLL STABILIZATION WITHOUT A FREE GYRO 


181 


to move the control flaps so that the restoring torque 
is directly proportional to the time. If aerodynamic 
friction is neglected, the total angular acceleration 
will be expressed by 

1$ = -kt (9) 

in which I is the moment of inertia about the roll 
axis. By integrating equation (9) 

4 > = ~ 2£ + 0o (.10) 

and, eliminating t between equations (9) and (10) 

* = - + <a° (id 

which gives the relation between the velocity and the 
acceleration. 


Control starts at A, and <£ decreases until <f> = <j> 0 /2 
at point B (Figure 22). The direction of motion of 
the control flaps is then reversed, and the correcting 
torque decreases in proportion to the time. At 0, 
<j> and </> reach zero simultaneously, and there is no 
overshoot. 

If the ailerons were not reversed at some arbitrary 
value of roll velocity shown here as <p 0 /2, the missile 
would continue to return to a stationary condition 
in the roll sense but would have a considerable dis¬ 
placement of the ailerons resulting in a large roll ac¬ 
celeration, as indicated in the extension of A B to U 
in Figure 22. While there is no fundamental method 
in mechanics to store acceleration, the structure of a 
missile with ailerons does exactly that because of the 
time required to move the control surfaces. Thus, 
while it would appear that a system sensitive only 



Figure 21. Fluctuation in output of wide-angle scanner and its effect on flap motion. 







182 


MISCELLANEOUS CONTROL SYSTEMS 


to velocity is fundamentally not oscillatory, oscilla¬ 
tions creep in by virtue of this apparent storage of 
acceleration. 

With reversal of the ailerons at the appropriate 
time, the roll velocity will reach zero at the instant 



Figure 22. Relationship between roll velocity and 
acceleration for nonoscillatory stabilization. No initial 
acceleration. 


faces reduce the total roll torque, and therefore the 
acceleration, to zero at D. At this point conditions 
are the same as at A in Figure 22 and the curve sub¬ 
sequently traced, DEO in Figure 23, will be similar 
to ABO in Figure 22. 

If, instead, 0 O and 0 O are opposite in sign, the course 
of events will be as indicated in Figure 24, starting 



Figure 24. Nonoscillatory roll stabilization with ini¬ 
tial roll velocity and acceleration of opposite sign. 


when the acceleration is also zero, and no further 
angular motion Avill take place. 

Case 2. Initial acceleration not zero. To eliminate 
one of the unreal assumptions, assume that asym¬ 
metry in the bomb construction introduces a roll 



Figure 23. Nonoscillatory roll stabilization with ini¬ 
tial roll acceleration. 


torque. This adds the constant term <?‘ 0 to equation 
(9). Graphically it is expressed directly by changing 
the starting point of the <j> vs <j> graph as in Figure 23. 
If both 0o and 0 O are positive, the curve will start 
at a r point such as C in Figure 23. The control sur¬ 


at F. Here two reversals of the ailerons, each at the 
instant when the angular velocity has been reduced 
to half its most recent maximum value, are necessary 
to stop rotation at an instant of zero acceleration. 

Case 3. Initial velocity high. If the initial roll ve¬ 
locity, 0o, is very high there may be oscillation. This 



Figure 25. Oscillatory stabilization produced by exces¬ 
sive initial angular roll velocity. 


oscillation derives from the memory circuit and can 
continue only if memory is permanent—that is, if 
direction of the aileron motion always reverses when 
0 = A:0 max . Consider a missile with initial conditions 
represented by A in Figure 25. 

The roll velocity will increase until the aileron ex¬ 
cursion has developed sufficient torque to kill the 





















ROLL STABILIZATION WITHOUT A FREE GYRO 


183 


initial roll acceleration. From this point B, the ve¬ 
locity will decrease with increasing (negative) ac¬ 
celeration until the* limit of aileron travel is reached 
at C. From C the velocity will decrease under con¬ 
stant acceleration until the critical velocity, say «, 

is reached at D. Here the direction of aileron travel 
reverses, and the velocity and deceleration decrease 
along the locus DEF. At F, on the same ordinate as 
D, reversal again takes place, and stable oscillation 
is set up about a mean value of roll velocity k<j > max . 

Memory is established by charge on a condenser. 
Since the time constant cannot be infinite, the volt¬ 
age on the condenser and, therefore, the critical value 
of velocity at which the ailerons reverse their direc¬ 
tion gradually decay toward zero. The locus DEFH, 
therefore, drifts slowly toward the origin. 

When a point J is reached such that the next swing 
will result in a reversal of velocity, damping will set 
in, since memory is erased with each crossing of the 
axis of roll acceleration. 

Case 4. Measured angular velocity limited. If, in¬ 
stead of always reversing the ailerons at the maximum 
value of <j>, a limit <j> m is set to the angular velocity 
that can be registered linearly by the control circuits, 
any increases in velocities higher than <j> m will be dis¬ 
regarded; that is, a voltage used to operate the con¬ 
trol would be proportional to (j> in the range from 
— <j> m to + <p m , and constant for values of <j> outside 
these limits. Then, in Figure 25, the actuating volt¬ 
age would be constant from A to U. There the con¬ 
trol would operate as before, the flaps would reverse 
at J, where the velocity is half that at U, and the 
bomb would come to rest as indicated. 

Evaluation of <t> m . From Figure 25 we can obtain 
information for design. If we define (j> r as that value 
of <j> at which the flaps must be reversed to stop the 
roll without oscillation ( SQT ), the value of <j> r , and 
hence <j> m (= 2 <j> r ) will be uniquely determined by the 
line CJ. This in turn is determined by T m , the torque 
generated when the flaps are in their position of maxi¬ 
mum deflection. 

If I is the moment of inertia about the longitudinal 
axis, X the coefficient of aerodynamic friction, and r 
the run-out time of the flaps, it can be shown that 



If in turn Xr/7 is small, this reduces to 

*/=§ (13) 


In this expression all parameters are determined by 
the construction of the missile, which permits the 
evaluation of <p r . 

Preferred Arrangement of Control System. It is im¬ 
plicit in the foregoing that <j> r should be made small 
enough so that the locus will reach the axis of <po on 
the first reversal of the ailerons. <j> m need not exceed 
2(j> r . 

When the value for <j> m has been chosen in this way 
the controlling voltage is made proportional to <f> for 
absolute velocities smaller than <p m , but the rate gyro 
is blocked so that higher values of <£ cannot be regis¬ 
tered. In this way sensitive control is provided for 
low velocities, and at the same time a means is avail¬ 
able to handle high velocities in the manner indi¬ 
cated in Figure 25. 


Effect of Time Lags 


Unavoidable time delays in operation of relays and 
aileron mechanisms must introduce residual oscilla¬ 
tion in any actual device. Assume fqr example that 
the representative point is traveling from F to A 
along the locus of Figure 26. Instead of reversing 
instantaneously at A, where <j> = }/ 2 <i>m, and proceed¬ 
ing to 0 along the dotted path, the mechanism suf¬ 
fers a time lag and does not reverse until some point 
B has been reached. The path BCE will be followed, 
reversal being called for at D and actually accom¬ 
plished at E. It is evident that sustained oscillation 
must occur. 

If T 0 is the total time delay in seconds, P the period 
of the residual oscillation, and </> m its amplitude, it 
can be shown that 

P = = 13Jr ° seconds ( 14 ) 


_ W2T m T<? 
0 ”* 3(\/2-1) 3 t7 


26.6 T m rf radiang n (15j 

T 


B See Experimental Investigations in Connection with High- 
Angle Dirigible Bombs—-The AZON Bomb , Gulf Research and 
Development Co., October 15, 1943. OSRD No. 3086, in 
which the author finds that the amplitude of oscillation is: 
(r) 2 (2fc - rYTm 
8 (k - r) 2 / 

where k is a constant relating the coupling between the rate 
gyro and the free gyro which were used in the stabilizing of 
Azon. In equation (15), reducing r to a very small value 
causes <£ max to increase. In the Gulf equation, making r equal 
to zero reduces the amplitude of oscillation,to zero. 

Actually the Azon ailerons were powered by solenoids and 
the action was made very rapid (r = 0.1 second). The ob¬ 
served oscillation during drops (see Chapter 2) was very 
small. to 









184 


MISCELLANEOUS CONTROL SYSTEMS 


It is evident that (1) r G should be minimized; (2) de¬ 
creasing T m and increasing r improve performance 
as far as oscillation is concerned. 


9.5.3 


Control Circuits 


Figure 27 shows the essential storage circuit which 
causes reversal of the ailerons when <j> is reduced to 
any desired fraction of <t> m - 



Figure 26 . 
anism. 


Oscillation caused by time delay in mech- 


Potentiometer 1 is driven by the rate gyros. A posi¬ 
tive voltage charges condenser 4 but is blocked from 
condenser 10 by diode 11. A negative voltage charges 
condenser 10 but not 4. The polarity is referred to 
terminal 8, at the midpoint of the potentiometer. If 
<j> increases positively, the charge on 4 builds up to 
a value E m , corresponding to <t> m . Then as 4> decreases, 
the time constants of the circuit are sufficiently long 
(of the order of 20 seconds) so that the charge on 4 
remains constant until <j> reaches zero. A relay (not 
shown) then short-circuits the condenser. Similarly 
a negative swing of <i> will impress a voltage on 10 
representative of 


The voltage E 2 across terminals 7 and 8 can be 
computed to be 


E 2 = 


E - KE V 
2 + K 
E - HE m 


( 16 ) 


If K = 2, H = V 2 ) if K = 1, H = When E 
drops below HE m , the polarity of E 2 reverses. This, 
through amplifiers, a polarized relay, and magnetic 
clutches, is made to operate the ailerons, 
where K = proportionality constant in resistors 6 
and 9 of Figure 27, 

H = h, 

<t> r 

E = voltage across potentiometer 1. 



9.5.4 


Experimental Results 


Using a model bomb in which the control appara¬ 
tus was mounted, wind-tunnel tests at low wind ve¬ 
locities were made. Direct measurements were made 
of T m , t, and the time for various amounts of rota¬ 
tion. The period of oscillation of the actual device as 
built was measured and found to be 1.37 seconds. The 
amplitude as measured and computed agreed well at 
about 16 degrees. 

Faulty adjustment of the potentiometers, so that 
reversal occurs at values of <j> greater than <j> m / 2, can 
be shown to induce sustained oscillation accompanied 
by drift. This was confirmed by test. 

By the time this work was completed it was evi¬ 
dent that the bomb was operable. Nevertheless, 
stabilization of the Gulf bomb had by then been ac¬ 
complished by other means and there was no further 
demand for this device. 























THE TO-AND-FRO SCANNER 


185 


The following comments are pertinent. 

1. If the reversal of the ailerons occurs too soon, 
because of improper functioning of the storage cir¬ 
cuit, roll oscillation superposed on a drift of roll 
orientation must ensue. 

2. Functioning of the storage circuit is contingent 
upon discharge of the memory condensers every time 
the angular velocity becomes zero. This is accom¬ 
plished by a circuit containing two d-c amplifiers and 
a relay. Sensitivity is a function of the adjustment 
of these amplifiers. 

3. The whole device is operable when properly ad¬ 
justed, but has many more parts, more adjustments, 
and hence more chance of malfunctioning than par¬ 
allel devices developed at Gulf and at Douglas. 

4. In retrospect it is difficult to see what major ad¬ 
vantage is obtained by this rather elaborate circuit 
over the very simple devices developed for the same 
purpose by MIT (see Chapter 2) and Douglas (see 
Chapter 4). This conclusion could not, of course, have 
been reached a priori. The contribution of this con¬ 
tractor to the analysis of roll-stability problems was 
valuable and encouraged the Division to continue 
the experimental work necessary to embody the prin¬ 
ciples of his analysis in a working system. 


96 THE TO-AND-FRO SCANNER 

Section 4.2.4 described how Roc was built so that 
the rate of change of transverse lift was proportional 
to the error heading and to its first derivative. All the 
early Gulf bombs (Section 2.2.7) also used propor¬ 
tional control until (Section 2.7.3) it was discovered 
that, for Azon at least, it was not needed. In an at¬ 
tempt to supplement the Roc and Felix projects by 
developing an infrared scanning device capable of 
giving quantitative measurement of error angle 
(rather than quadrant indication alone), Electro- 
Mechanical Research, Inc., was directed to continue 
portions of the work they had started independently 
on a heat-homing bomb similar morphologically to 
Roc. 

Under the EMR contract, a roll-stabilizer was de¬ 
veloped which needed no free gyro (Section 9.5), 
bolometers were improved (Section 9.6.2), six suc¬ 
cessively refined scanning devices using to-and-fro 
scanning were built, and contributions were made to 
the theory of bolometer response to modulated heat 
energy. No way was found by this contractor, how¬ 
ever, to achieve proportional control. 


961 Horizon Difficulties 

A heat-homing device using a conical scan, such as 
Felix and Dove Eye (Chapter 3), cannot be allowed 
to see the horizon. Since such a device operates on 
the largest thermal discontinuity within its view, it 
tends to attempt the maneuver of homing on the 
horizon. On the other hand, if the area scanned can 
be kept wide horizontally but narrow vertically, it is 
possible to prevent the scanning element from seeing 
the horizon even if the angle of flight is fairly flat, 
say 11 degrees. The only method of obtaining a scan¬ 
ning pattern of this shape utilizes a to-and-fro 
scanner. 

9 6 2 S-3 Scanner 5 

The first laboratory model of such a scanner was 
equipped with lights to indicate right or left target 
bearing, and operating relays in lieu of bomb controls. 
An 8-in. parabolic mirror (focal length 6 in.) reflected 
the target image upon a bolometer strip. The optical 
system, including the thermal receptor, was oscil¬ 
lated at 12 c about a vertical axis through the apex 
of the mirror. A narrow-band amplifier, tuned to the 
scanning frequency, amplified the bolometer output. 

Theory of Operation 6 

Consider a parabolic mirror which oscillates in a 
to-and-fro manner so that it scans a field of view con¬ 
taining a thermal target. If an ideal bolometer, i.e., 
one having a very small mass and specific heat, is 
located at the focus of such a mirror, its resistance 
will rise sharply each time the image of the target 


BOLOMETER 



Figure 28. Typical bolometer input circuit. 


falls upon it. If the bolometer is connected in a bridge 
circuit such as that in Figure 28, the rise in resistance 
will be accompanied by a voltage pulse at the second¬ 
ary of the transformer. 

In general there will be two voltage pulses per 
cycle of scan. If the target is dead ahead, the pulses 







186 


MISCELLANEOUS CONTROL SYSTEMS 


will be evenly spaced at double the scanning fre¬ 
quency. At some error angle 6, the pulses will be 
spaced at intervals of tt ± 26, where 2tt is the period 
of the scanning cycle (Figure 29). In the special case 
where the target is at the edge of'the field of view, 


p -77-+2S - ^77-29 p- 

77+20 - p 

1 

| 

n ! n n j 

n 

n | 

1 

f-i — 1 SCANNING CYCLE - 


— i — 

i 

i 

L- 

- 277 - 

i 

-H 


' i ' 

Figure 29. Pulses due to thermal target. 

6 = 7t/ 2 and tt — 26 becomes zero, so that the two 
pulses coalesce. 

Now let the secondary of the transformer in Figure 
28 supply an amplifier tuned to the scanning fre¬ 
quency, the frequency at which the mirror oscillates, 
but with sufficient band-pass to prevent material 
phase shift. The output of this amplifier will consist 
principally of the fundamental components of the 
Fourier expansion of the pulses at scanning fre¬ 
quency. Each fundamental component will be in 
phase with its pulse. 

If a rotating vector so represents the position of 
the axis of the mirror during a scanning cycle, the 
diagram of Figure 30 reproduces in a polar plot events 
during the scanning cycle in Figure 29. At time 0 the 


the system is in an extreme position and starts back. 
When it is short of the mid-position r by the error 
angle 6, it will again pick up the target, producing a 
second pulse with a fundamental frequency com¬ 
ponent V\. The sum of these two sine waves at scan¬ 
ning frequency is 

E = 2F lS in^- (17) 

"t7max 

The magnitude of the voltage produced as a func¬ 
tion of error angle can now be plotted from equation 
(17). It must be emphasized, however, that this plot 
(Figure 31) does not represent the instantaneous out¬ 
put of the amplifier as the scanning mirror makes a 
swing through half a scanning cycle. Rather it is a 
steady-state plot of the magnitude of a sinusoidal 
voltage which varies with error angle. 

As the voltage passes through zero, corresponding 
to zero error, there is a phase reversal. With a left 
error the crest of the voltage occurs when the system 



ERROR ANGLE 

Figure 31. Output of S-3 scanner preamplifier. 


TT 



37T 

2 


Figure 30. Vector representation of fundamental 
component of pulses from target. 

scanning system is looking dead ahead. At some later 
instant, after the system has swept through the error 
angle 6, the receptor will pick up the thermal target 
and produce a voltage pulse whose fundamental com¬ 
ponent is in phase with the pulse at V\. At time 7t/2 


is at its extreme of left travel, that is, at time — tt / 2 ; 
with a right error the voltage is at a maximum when 
the system is at its extreme right travel, that is at 
time 7 t/2. 

Sense of direction is readily obtained if a syn¬ 
chronous voltage is added to the output of the pre¬ 
amplifier at a phase to support the output when the 
error is in one sense and to suppress it if the error is 
in the reverse sense. Equation (17) represents the 
maximum value of the fundamental component of 
the output. Its instantaneous value is 

ei = 2Vi sin sin (ut ± ^ (18) 

The double sign for the phase angle indicates the 
phase reversal which takes place as 6, measured from 
the scanning axis, changes sign. The added voltage 
can be written as: 

e 2 = E sin -f- (19) 

In this case the sign of the phase angle is arbitrary 
but not ambiguous. The sum of these two voltages 














THE TO-AND-FRO SCANNER 


187 


as a function of 0 is plotted in Figure 32. It is seen 
to have the form 

E + 2V l siti^- (20) 

""max 

The curves of Figures 31 and 32 are plotted as pure 
single-frequency functions since all the development 
of the theory has been concerned with the funda¬ 
mental component of the detected pulse. The addi¬ 
tion of harmonics distorts the shape of the curve but 



ERROR ANGLE 

Figure 32. Combined output of preamplifier and 
synchronous local voltage. 

does not introduce modes unless the harmonics are 
high compared to the fundamental (Figure 33). 

The local voltage was produced by a synchronous 
generator driven from the same motor which oscil¬ 
lated the scanning mirror (Figure 34). The combined 
outputs of the preamplifier and the generator were 
amplified in a narrow-pass pentode stage and sup¬ 
plied to a two-channel rectifier and amplifier, which 
operated the control relays. 

The Bolometer 

The Division had three contractors working on 
bolometer development. Felix used nickel strips, 


Dove Eye used thermistors, and the EMR to-and-fro 
scanner used metallic gold evaporated upon thin films 
of cellulose nitrate. Each of these bolometers had its 
own characteristics, but all had a useful sensitivity 
of the same order of magnitude, 2.0 X 10 -8 watt per 
sq cm. Some experimental models, both of Felix and 
of Dove Eye, used film-type bolometers. 

Early EMR bolometers consisted of a thin strip of 
gold evaporated upon a cellulose nitrate film (0.1 



Figure 33. Typical signal from S-3 scanner as func¬ 
tion of error angle. 


micron thick) with zinc black (later replaced by gold 
black) evaporated in turn upon that. Sodium chloride 
windows, ground flat and sealed with beeswax, al¬ 
lowed transmission of infrared energy to the bolom¬ 
eter. After a short time the seal usually leaked and 
the vacuum was destroyed. 

As development proceeded, glass cases .were re¬ 
placed by successively smaller metal cases, various 
protective coatings for the salt windows proved none 




Figure 34. Block diagram of scanner. 





































188 


MISCELLANEOUS CONTROL SYSTEMS 


too good and silver chloride windows were substi¬ 
tuted, and finally gold blackening in the presence of 
traces of tellurium was shown to give the most satis¬ 
factory performance. 

A lengthy study of desirable characteristics of bo¬ 
lometers 7 showed that a metal plate about 0.006 in. 
behind the film should give greater sensitivity in 
nitrogen (100-mm pressure) than under a vacuum; 
this was confirmed. Of the metals tried gold proved 
best. 

Sensitivities measured were of the order of 1 v per 
watt per sq cm at the grid of the first tube, in terms 
of the total energy field at the bolometer. It is per¬ 
tinent at this point to mention that MIT, Polaroid, 
and EMR all measured sensitivities of bolometers by 
slightly different methods and obtained results which 
agreed within about 10 per cent. Polaroid has shown 
that, because of variation in the absorption coeffi¬ 
cient of the blackening resulting in spectral distribu¬ 
tion of the incident energy, this much difference is to 
be expected. For a given total quantity of energy in¬ 
cident upon the bolometer, the amount absorbed 
(and hence effective) from a short-time high-temper¬ 
ature pulse will differ from that from a longer low- 
temperature pulse. Therefore, the measured sensi¬ 
tivities will differ in the two cases. No general law has 
been adduced, and the only remedy is to specify ex¬ 
perimental conditions exactly. Polaroid has also 
shown that the time constant of the entire system is 
an important variable factor in the measurement of 
sensitivity. 

Results 

When tested at MIT the scanner was set up look¬ 
ing vertically downward at a tank of water 430 cm 
(14 ft, 2 in.) below. The blackened end of a brass cyl¬ 
inder, 15 sq cm in area, was the target. A difference 
of 1.5 C to 2 C between target and surroundings, cor¬ 
responding to a signal intensity of 0.02 microwatt 
per sq cm, operated the signal lights in correct order 
as the target moved across the field. 

Taken outdoors and fitted with a magnesium oxide 
filter to eliminate visible light, the device was tested 
on ships in Boston Harbor. Correct indication was 
obtained from a destroyer at 3,100 yd but not at 
4,500 yd. 

Operation was seriously hampered by the necessity 
for frequent readjustment of the level of the local 
signal. This was caused by background irregularities, 
which of course served as false targets. 


Conclusions 

The device as built was much too large and cum¬ 
bersome; it possessed moderate sensitivity; it was 
operable only when adjusted frequently to eliminate 
the effect of variation of background noise. It was 
therefore not suitable as a military device. 

9 6 4 S-4 and S-5 Scanners 

The S-4 was essentially the same type as the S-3 
except that the frequency and amplitude of the scan 
were adjustable. Amplitude was variable from 0 C 
to ±8 C, frequency from 3 to 12 c. Work with the 
S-3 had indicated that there might be advantages to 
using higher harmonics of the fundamental scanning 
frequency, so two amplifiers were used simultaneous¬ 
ly in series with the bolometer, one tuned to 12 c and 
the other to 24 c. 

The S-5 scanner was designed for compactness. A 
spherical mirror with a diameter of 3^ in. and a focal 
length of 2.25 in. collected the incoming energy. The 
bolometer was mounted in a hole in the center of 
this mirror. Energy collected by the spherical mirror 
was reflected by it upon a plane mirror 1J4 in. in 
diameter, which in turn reflected it upon the bolom¬ 
eter. The flat mirror was oscillated about a vertical 
axis. The advantages of this type of construction 
were (1) small power consumption and (2) compact¬ 
ness. Its main fault was that it was useful only for 
narrow scanning angles. When the scanning angle 
was such that the field of the flat mirror passed the 
edge of the curved mirror, energy from a uniform 
background would be modulated at the scanning fre¬ 
quency, giving a spurious signal. This limitation re¬ 
moved the S-5 from serious consideration. 

Harmonic Scanning 

From this point on the discussion will deal largely 
with operation based upon use of the second har¬ 
monic of the scanning frequency. It is necessary, 
therefore, to see what is to be gained therefrom, and 
at what cost. 

If the scanner consists of a bolometer and a fixed 
mirror, the plot of output vs azimuth angle will be 
qualitatively as shown in Figure 35A. 8 

Now if an oscillating mirror is used, and the bo¬ 
lometer output is fed to a sharply tuned amplifier 
responsive only to the mirror frequency, the plot of 
output vs angle will be as in Figure 35B. The exact 




THE TO-AND-FRO SCANNER 


189 


shape depends upon the design factors, but it will 
always be of the general shape shown. 

If the amplifier is tuned to the second harmonic, 
twice the mirror frequency, the response curve will 
be as shown in Figure 35C. This curve will obtain if 
the scanning angle and the bolometer strip are both 
narrow. For larger angles, the relative heights of the 
center peaks and the outer peaks will change. 

Figures 35D and 35E show the patterns obtained 
with the third and fourth harmonics respectively. It 
is evident that the number of peaks is n + 1, if n 
is the order of the harmonic used, and the number of 
interior minima is equal to n. 

/\ | /n 

A B 


rS\ J\\ ^rJ\j\r, 


Figure 35. Output of tuned amplifier with error angle: 
(A) fixed mirror, (B) amplifier tuned to scanning fre¬ 
quency, (C) amplifier tuned to two times scanning 
frequency, (D) amplifier tuned to three times scanning 
frequency, (E) amplifier tuned to four times scanning 
frequency. 


Fundamental Response 

This results from the general analysis described in 
Section 9.6.2. Two pulses per cycle are produced by 
the bolometer. The amount of the voltage output de¬ 
pends upon the position of the target with reference 
to the swing. With the target at either end of the 
scanning arc, the fundamental-frequency outputs add 
up to a maximum. As the target nears the center, the 
vectors representing the two pulses rotate in opposite 
direction, reaching direct opposition with the target 
on course. The two pulses being equal and 180 de¬ 
grees out of phase, a minimum voltage output is ob¬ 
tained. 


Higher Harmonics 

Each series of pulses will also generate harmonic 
components of voltage in the output, and each pair 
of such components of a given frequency can be rep¬ 
resented by a pair of vectors, as shown above. With 
the target at the extreme right of the swing, a maxi¬ 
mum is obtained as before. As the target moves to 
the left, however, the phase shifts will be propor¬ 
tional to the order of the harmonics. For the second 


harmonic the two pulses will be 180 degrees out of 
phase, producing a minimum at two points along the 
scanning cycle. Similar analysis applies to higher 
harmonics. 

When the individual energy pulses are short (nar¬ 
row image, narrow bolometer, wide scanning angle), 
the amplitudes of the fundamental and of the lower 
harmonics are approximately equal. This is based 
upon the following reasoning. 

In the case of rapidly interrupted energy of rec¬ 
tangular waveform, the amplitude of the ?ith har¬ 
monic of the output voltage is 

v _ 22? • /nu)At\ /01X 

En n*V 1 + U n 2 Sm ( 2 ) (21) 

where E = voltage when ample time is allowed; 

E n = voltage of nth harmonic; 
n = order of harmonic; 
co/ 27r = # frequency of pulse; 

At = duration of pulse. 

Un is analogous to the ratio of reactance to resist¬ 
ance in an electric circuit containing inductance and 
resistance in series. In the simple case of a bolometer 
which loses heat only by radiation 


U n 


uojH 

4crT\ 3 


( 22 ) 


In this a — effective radiation constant of the strip; 
H = total thermal capacity of the strip; 

T i = absolute temperature of the strip. 


Since U n is directly proportional to the frequency 
of interruption of the received energy, equation (21) 
indicates that the response should be proportional 
to l/\/l + k 2 f 2 . This is substantially true experi¬ 
mentally. 

Other things being equal, equation (21) shows that 
the amplitude of the nth harmonic of the bolometer 
output is proportional to 1/n sin (nwAt/2). For small 
values of the quantity in parentheses, this reduces to 
uAt/2, and the amplitudes of the lower harmonics 
are equal. 

The above has been predicated upon a rectangular 
pulse. Theory and experiment agree that the value of 
the peak amplitude will be less if the pulse is a half 
sine wave instead of rectangular, but the relation¬ 
ship among the harmonics is undisturbed. 

So far we have deduced that the target signal 
should be approximately the same amplitude whether 
the amplifier is tuned to the fundamental or to one 
of the lower-order harmonics. Similar analysis shows 
that slow changes in the intensity of a uniform back- 














190 


MISCELLANEOUS CONTROL SYSTEMS 


ground, or a uniform gradient across the background, 
should introduce greater background noise into the 
fundamental than into each successive harmonic. 
This was also verified experimentally. 

Choice of Second Harmonic 

If the fundamental or the third harmonic is used, 
distinction between right and left indication is by 
phase determination, as has been seen in the S-3. 
With the third harmonic, however, plaks 1 and 4 
(Figure 35D) have the same phase as peaks 3 and 2 
respectively, and if the target is at the extreme of the 
field of view, the resulting voltage will be of the 
wrong phase and the control reversed in sense. This 
limits the use of the third harmonic to targets near 
the center of the field—obviously impossible to 
guarantee in a military application. 

Using the second harmonic—which has been seen 
to give a better signal-to-noise ratio than the funda¬ 
mental—the response curve (Figure 35C) is symmet¬ 
rical, and the phase of the voltage cannot be used for 
control. In the case of the bomb control, this led to 
the design of the S-8 scanner, in which two halves of 
the mirror oscillated independently, and right-left in¬ 
dication was derived from this fact. Before that was 
built, however, much experimental evidence verify¬ 
ing these theoretical considerations was acquired in 
work on scanners S-4, S-6, and S-7. 

Laboratory Results 

In the laboratory, with a target showing a uniform 
gradient across the field of view, the second-harmonic 
noise for a scanning angle of 3 degrees was shown to 
be from V 4 to Vi 0 the fundamental-frequency noise. 

In the field, at an observation station overlooking 
Galveston Bay, results were similar. Recording the 
output voltages of the two amplifiers simultaneously 
by means of recording voltmeters, curves similar to 
Figure 36 were obtained. These typical runs show 
that differentiation between the target and the back¬ 
ground noise is very much better with 24 c than with 
12 c (the fundamental). 

9 6 6 S-6 Scanner 

For use in a target survey to be made by a Navy 
Bureau of Ordnance plane, a scanner was built to 
operate on both 12 and 24 c, as before, without right- 
left indication. Essentially it was very similar to the 


S-4 except that it was built for mounting in the plane 
and therefore was not geared to the recording volt¬ 
meters. 

At the wide scanning angle of 14 degrees, the ad¬ 
vantages of second-harmonic tuning were not very 
great, and results in general were none too satisfac¬ 
tory. Among the reasons were: 

1. The scanner was being compared by the Navy 
with instruments having much smaller fields of view. 
They necessarily showed less background noise (and 
consequently higher signal-to-noise ratio). 

2. It was a laboratory device operated by field 
personnel. Such a combination frequently leads to a 
feeling by the operators that the apparatus is un¬ 
reliable, by the development personnel that it is in¬ 
competently used. 

3. Sensitivity (again a function of the scanning an¬ 
gle) was too low, so that at desired ranges unreliable 
target indication was observed. 

It is important that liaison on this project was not 
adequate to keep either the Division or the contractor 
fully advised as to the success or failure of the scan¬ 
ner in the field. As has been repeatedly pointed out, 
only by continuous correlation of all the data taken 
by all agencies associated with a project can success¬ 
ful work be done. 

9 6 7 S-7 Scanner 

While the S-6 was on loan to the Navy, a similar 
scanner was built for laboratory and field use. 

• I 

Construction 

This is shown in Figures 37 and 38. 

The new features were provision (1) for adjust¬ 
ment of the scanning angle (from 0 to ± 1.75 degrees) 
by the larger knob seen at the back of the case, and 
(2) for focusing the bolometer during operation. 

The reduction in size from the S-3 is evident. The 
desirable reduction in electrical complexity is also ap¬ 
parent. Target indication was accomplished by lights 
operated by a relay connected to the output of the 
6C8 tube seen at the right. 

When the scanner is pointed directly at the target, 
the phase of the output voltage of the amplifier can 
be adjusted (by the phase-shifting network at the top 
left) so that one element of the 24-c synchronous rec¬ 
tifier produces a maximum positive potential. The 
other element of the rectifier is permanently 180 de¬ 
grees out of phase with the first; it therefore develops 



THE TO-AND-FRO SCANNER 


191 





Figure 36. Test results with to-and-fro scanner. 


a negative output potential, and the two sections of 
the 6C8 twin triode are acted upon oppositely by the 
voltages from the two rectifiers. If each plate is con¬ 
nected in series with one winding of a differential 
relay, the armature will be drawn to one side. 

Now if the scanner is turned so as to bring one of 


the side peaks of the response pattern into play, the 
phase of the amplifier output will be reversed. This 
reverses the polarity of the d-c potential upon the 
6C8 grids and reverses the relay. Signal lamps con¬ 
nected to the relay can then show whether the target 
is in the center or a side lobe. 


























































































































































































































































































































































192 


MISCELLANEOUS CONTROL SYSTEMS 


Results 

Field runs (Figures 39 and 40) showed the typical 
two-peak pattern for the 12-c log and three peaks for 
the 24-c, together with lowered background noise for 
the latter. 

The major conclusions from the experience with 
scanners S-3 through S-7 were that second harmonic 


tuning gave a greatly improved performance, that 
the scanning angle must be at a minimum, and that 
with proper adjustment a signal strength below 1 
microwatt per sq cm at the mirror was adequate for 
reliable operation. The first two conclusions, incon¬ 
sistent with the principles of the S-3 (which involved 
wide-angle scanning and phase discrimination for 
right-left indication), forced design changes in S-8. 



Figure 37. Scanner head (top) and complete S-7 scanner (bottom). 








THE TO-AND-FRO SCANNER 


193 



S-8 Scanner 


Construction 


The foregoing portions of this section report the 
development of a principle for scanning which might 
be applicable to a missile. None of the scanners so far 
discussed was contemplated for such application. The 


Figure 41 shows views of the scanner as assembled 
and with the case removed. The first novel feature 
is the split mirror. The same type of spherical mirror 
is used as in previous models, but it is cut along a 



Figure 39. Field results with S-7 scanner (run from 
east toward north). 

S-8 scanner was an attempt to show the applicability 
of the principles discovered in development of the 
earlier models to the guidance of a glide bomb in 
azimuth. Control in the range sense was to be con¬ 
sidered later. After the termination of the project by 
the Division, the contractor completed this latter 
phase of the investigation under contract with the 
Air Technical Service Command. 



Figure 40. Field results with S-7 scanner (run from 
east toward south). 

horizontal diameter so that the two halves are, in ef¬ 
fect, separate mirrors which can look in different 
directions. The two halves are oscillated at the same 
frequency, but both the angle between their axes and 
the relative phase of the motion are continuously ad- 



















































































































194 


MISCELLANEOUS CONTROL SYSTEMS 


just able by knobs. This latter makes it possible to 
distinguish between the signals derived from the two 
halves by using phase-selecting circuits. 

Secondly, the scanner is mounted on a motor- 
driven turntable to permit automatic tracking of the 
target. 

Thirdly, driven by the scanning motor are two 2- 
element 24-c synchronous rectifiers whose function¬ 
ing is described below. 

Amplifier Circuits 

The amplifier circuit (Figure 42) consists of four 
parts: a tuned preamplifier, a volume limiter, the rec¬ 
tifier and d-c output circuits, and a vacuum-tube 
voltmeter incorporated for experimental convenience. 

It will be noted that the wide-band amplifier of the 
S-3 has been eliminated. Instead, the phase-shifting 
network at the upper left adjusts for any phase shift 
through the tuned amplifier. 

Following the phase shifter comes a three-stage 
limiter. This simple circuit will handle a very wide 


range of input voltage with practically no change in 
output level. 

Voltage from the last limiter tube is supplied to the 
four rectifier elements. The two upper rectifier ele¬ 
ments deliver output voltage of opposite polarity. 
The two lower elements are also 180 degrees out of 
phase with each other, but are in quadrature with the 
lower elements. 

Synchronous Rectifiers 

These rectifiers serve two functions: first, they pro¬ 
vide a means for distinguishing between the signals 
received from the two halves of the mirror, and sec¬ 
ond, they enable all three lobes of the response pat¬ 
tern of one mirror to operate the relay in the same 
way. Since the signal from the center lobe is out of 
phase with that from either of the outer lobes, two 
rectifiers 180 degrees apart must be associated with 
each mirror. One element develops a positive poten¬ 
tial for closing the relay when the target is in the cen¬ 
tral lobe of the pattern, while the other develops the 




Figure 41. S-8 scanner. 


















0.5 MEG 0.5 MEG 


THE TO-AND-FRO SCANNER 


195 



Figure 42. Wiring diagram of S-8 scanner. 



















































































































































196 


MISCELLANEOUS CONTROL SYSTEMS 


necessary positive potential when the target lies 
within either of the outer lobes. 

The magnitude of the output of a synchronous rec¬ 
tifier is dependent upon the phase of the impressed 
a-c potential with respect to the rectifier. When the 
phase difference is 90 degrees, the output is zero. The 
two pairs of rectifiers shown in Figure 42 are in 
quadrature with each other. The two halves of the 
mirror are 45 degrees out of phase, which puts the 
second-harmonic signals from the two mirrors also in 
quadrature. Therefore, the upper pair of rectifiers 
will respond only to signals from the right-hand mir¬ 
ror (one looking to the right) and the lower pair only 
to signals from the left-hand mirror. 

The differential relay shown operates two indicat¬ 
ing lights and also two single relays for operating the 
turntable motor. 

Results 

In the laboratory, while the amount of hunting de¬ 
pended upon the strength of the target, it was of the 
order of ±1 to 2 degrees. The minimum radiation 
field strength to operate the steering mechanism un¬ 
der favorable conditions was 0.01 microwatt per sq 
cm at the mirror. This was when the scanning angle 
was ± 1 degree and the angle between the two mirrors 
was 4 degrees, giving a total field of view of the bo¬ 
lometer of between 7 and 8 degrees. 

Outdoors, operation was irregular, some of the few 
possible targets observed being tracked reliably, oth¬ 
ers not. One successful run involved tracking a black 
barge and tug to a range of 2,850 yd. At 2,700 yd the 
Farrand Heat Meter recorded a field strength of 
0.11 microwatt per sq cm. On another run, a freighter 
was tracked at about 7,000 yd when its field strength 
was 0.03 microwatt per sq cm. 

Conclusions 

The S-8 showed much more satisfactory sensitivity 
than its predecessors. The signal-to-noise ratio is a 
function of the scanning angle, since the greater the 
angle the greater must necessarily be the probability 
of picking up background irregularity. Two mirrors 
at an angle with each other, each oscillated inde¬ 
pendently, will cover an effective area slightly larger 
than the sum of the angles measured (because of the 
length of the bolometer strip) with only a small 
scanning angle. The S-8 provided a reliable means of 
steering, the first in the series so equipped. 


Its sensitivity was of the order of magnitude ob¬ 
served for Dove Eye and for Felix—in other words, 
it should have been capable of operating bomb con¬ 
trols at usable distances. 

It had fewer tubes than its predecessors, and far 
less complexity than Dove Eye, but some of the ad¬ 
justments, for example the phase-shifting network, 
were critical. 


97 PROJECT BEETLE 9 

The difficulties enumerated above in connection 
with the quadrant photocell target seeker and with 
the wide-angle photoelectric scanner led the Division 
contractor, Bendix Aviation Ltd., to suggest the use 
of CW radio target seeking. In the centimeter range, 
this could have been the Moth technique (see Chap¬ 
ter 1) and would have been parallel with the testing 
technique used in the early phases of the radar¬ 
homing glide-bomb program. The technique envis¬ 
aged by the contractors, however, involved less com¬ 
plicated techniques than those of centimeter radar. 
The frequency proposed was in the neighborhood of 
200 me and the corresponding techniques were par¬ 
allel with those well established for radio goniometry. 
The Division supported this suggestion of its con¬ 
tractor and extended the contract (OEMsr-1002) to 
include the development of radio-homing devices for 
experimental purposes. 

9,7,1 Design of System 

The system proposed by Bendix consisted of four 
antennas symmetrically and orthogonally arranged 
about the roll axis of the missile near its nose. Each 
antenna was successively connected through a com¬ 
mutator to a fixed-tuned receiver. The signals from 
diagonally opposite antennas, suitably rectified and 
compared, gave an error signal proportional to the 
error in heading. 

The four antennas consisted of quarter-wavelength 
spikes radiating from a tapered nose on the forebody 
of the missile. A superregenerative detector was ca- 
pacitively coupled to each of the antenna spikes in 
rotation through a motor-driven commutator. A syn¬ 
chronously driven commutator at the output of the 
detector and amplifier distributed the rectified and 
amplified output to a smoothing filter, so that a 
smooth d-c voltage appeared across the output ter¬ 
minals of the filter. (See Figure 43.) 



PROJECT BEETLE 


197 



DRIVEN BY THE SAME MOTOR 


WAVE ANALYSIS 




OUTPUT OF DETECTOR AND AMP 



AFTER COMMUTATION 



UTILIZING 1 AND 3 AVERAGE D-C VOLTAGE 


Figure 43. Block diagram of Bendix Beetle system. 























































































198 


MISCELLANEOUS CONTROL SYSTEMS 


With the missile directly on course, the signal in¬ 
tensities at each of two diametrically opposite an¬ 
tenna spikes would be equal. With an error in course 
heading, the differential in signal strength at two 
diagonally opposite antenna spikes, shown in Figure 
43A, would be measured by the receiver, the output 
of which during a cycle of commutation would be as 



Figure 44. Beetle CW radio target seeker. 


shown in Figure 43B. Figure 43B shows the result of 
commutation and the appearance at the output of 
the filter of a smooth d-c voltage at the filter output 
terminals, proportional to the difference in received 
signals at the antenna. 

9,7,2 Tests with Beetle 

An experimental model (Figure 44) of this homing 
device was constructed, and was tested on the ground 


at Rosamond Dry Lake. The ground test showed 
reasonable promise for this system of control in the 
azimuth sense, although signals with azimuth errors 
in one direction were not precisely symmetrical with 
those in the reverse direction (Figure 45). 

The ground tests were followed by airborne passage 
tests to give an indication of its probable effective¬ 
ness in the range direction. Reflections from the 
ground gave erratic signals in the region adjacent to 
zero error heading. Since precise quantitative results 
are required in this region for accurate homing flight, 
this failure was considered sufficiently serious to 
justify shelving the project. It is probable that a 
more elaborate transmitting antenna structure than 
the one shown in Figure 46, with a suitable reflector 
behind the radiating element might have eliminated 
this difficulty. The appearance of satisfactory quad¬ 
rant photocells from Farnsworth Radio and Tele¬ 
vision Corporation, however, eliminated the pressing 
need for this type of target seeker and the project 
was dropped. 

9 8 ORGANIC TARGET SEEKING 101112 
9,8,1 General 

The simplest method of obtaining target discrimi¬ 
nation is through its recognition by intelligence. The 
Japanese suicide missiles employed exactly this tech¬ 
nique, using human organisms to guide the missiles 
into impact. While this technique was extremely ef¬ 
fective and imposed serious losses on our naval units, 
many misses were scored, and it was obvious after 
the program had been in operation for some months 
that the problem of steering an aircraft into a small 
target at combat speed is not a simple one. 



Figure 45. Error of signal from CW radio target seeker. 














ORGANIC TARGET SEEKING 


199 


Economy as well as considerations of humanity 
suggests the use of lower organisms to perform the 
functions of the Japanese suicide pilots. The use of 
trained house cats to steer a homing torpedo was sug¬ 
gested in World War I. In this program the sugges¬ 
tion was to use any easily available and readily 
trained animal organism. The contractor, General 
Mills, Inc., retained a psychologist in the field of 
organic behavior who made a broad preliminary 
study before submitting his proposal to NDRC. In 
his preliminary investigation the contractor con¬ 
cluded that the common pigeon or dove, columba 
livia, is a satisfactory organism for rapid and easy 
training and that the supply is considerably in excess 
of what would be required for military operations. 
Following the contractor’s initial survey, the Divi¬ 
sion made preliminary studies on the assumption that 
a satisfactory servo link could be developed to couple 
the response of the pigeon with the control surfaces 
of the missile. These studies justified a small con¬ 
tract (OEMsr-1068) for further developmental work. 

The advice of the Applied Psychology Panel of 
NDRC was solicited and their approval of the stand¬ 
ing of the psychologist was obtained. 

9 8 2 Training of Pigeons 

The training program is divided into three phases. 
In the basic phase, the birds are taught to peck at a 
distinctive object, not necessarily related to any mili¬ 
tary target. During this period, birds with apparently 
low intelligence levels and temperamental birds are 
weeded out. The basic training course consists of 
feeding the birds through a metal plate provided with 
a hole of distinctive shape. The entire plate is covered 
with a translucent membrane such as vellum paper, 
and the whole assembly illuminated from the rear. 
The birds learn to peck at the illuminated spot and 
recover grain located behind it. Additional feeding is 
restricted during this period to a level at which the 
birds lose weight. Their weight is plotted, and those 
birds which lose weight without developing a condi¬ 
tioned reflex to peck at an illuminated spot on a sheet 
of paper are culled out. 

The selected birds are then given a more advanced 
training in which they learn not only to peck at a dis¬ 
tinctive object but to peck more accurately and at a 
higher frequency. Accuracy is taught by reducing the 
size of the target which will result in delivery of food 
to reinforce the conditioned reflex. Higher frequency 
is taught by introducing into the training system a 


counting relay system which will deliver food only if 
a predetermined number of pecking impulses have 
been made against the processed plate within a pre¬ 
determined interval. Initially, a single peck will de¬ 
liver a few grains. At a more advanced stage, several 
pecks, for example five delivered within ten seconds, 
will result in the delivery of food. In the final stages 
of this phase of training, thirty pecks must be made 
in ten seconds in order for the conditioned reflex to 
be reinforced. 

This method of training builds up what is known 
as a reserve of impulses. If the bird is placed in the 
environment in which he has been conditioned to ex- 



Figure 46. Radio transmitter, including batteries. 


pect food by pecking, he will proceed to peck and 
will continue to peck although no food appears. It is 
this property which made the organism attractive for 
combat use. A complicated mechanism in the bomb 
to assure that the bird would be fed during the com¬ 
bat flight proved unnecessary. Once this reserve has 
been exhausted, however, there is doubt whether it 
can be restored. That is, if the bird’s faith is broken 
that pecking in the approved manner will produce 
food, he cannot be so trained as to restore that faith. 
Care must be taken, therefore, during the training 
program to see that at no stage are the birds sub¬ 
jected to a requirement which will extinguish the re¬ 
serve of impulses. 

After the pigeons have been trained to peck at dis¬ 
tinctive objects, the final phase in the program is to 
train them to identify and to peck at a specific, dis- 






200 


MISCELLANEOUS CONTROL SYSTEMS 


VEHICLE “ON TARGET,"cONTROLS AT"ZERO" 



vehicle’off target.'controls atcorrection" 


4 * 

TARGET 



Figure 47. Diagram of pneumatic servo for pigeon-controlled bomb. 


Crete object. Considerable work was done by the con¬ 
tractors from aerial photographs of ships at sea. Well- 
trained birds, approximately the top 50 per cent of 
the class, would invariably, at the close of the train¬ 
ing program, peck at the largest boat in a convoy 
irrespective of its position in the array. Other birds 
were trained on aerial photographs of cities, and were 
trained to select a particular building or street inter¬ 
section and would attack the image of this spot, ir¬ 
respective of the orientation of the photograph as 
presented to them. 

The basic training course, involving culling out un¬ 
satisfactory birds and training the remainder to peck 
at a satisfactory frequency at any distinctive object, 
occupied approximately six weeks. Having completed 
this course, the birds require approximately 48 hours 
briefing on a specific target. 

9,8,3 Servo Links 

This method of target seeking was intended for use 
with the glide bomb. The reason for selecting the 
glide bomb was that it flies with very little change in 
angle of attack. Therefore, the problem of the bird 
was much simplified. If he could keep the glide bomb 
pointed at the target, the probabilities of a hit were 
high. 


The scanning system consisted of a camera obscura 
enclosing the bird. A simple convex lens projected 
the image of the target field onto a sheet of ground 
lucite in front of the bird. The focal length and angle 
of the lens were selected to give a reasonably large 
field of view so that for any expected launching error 
the target would lie somewhere within it. 

The lucite screen which presented to the pigeon 
the field of view with the target on it was mounted 
on a double gimbal system. With an error in heading 
which would make the target appear up and to the 
right, the pecking of the pigeon would affect the 
gimbal-mounted screen to give corrective impulses 
downward and to the left. A pneumatic control sys¬ 
tem (Figure 47) was developed to connect the deflec¬ 
tion of the screen with the gyro pilot (see Section 1.6) 
of the glide bomb. There was considerable doubt as 
to whether this system of control was adequately 
quantitative to control a glide bomb without serious 
hunting. The fundamental problem in the servo sys¬ 
tem was to obtain an error signal proportional to the 
error in heading. The initial link in the chain which 
developed the signal produced an impulsive force of 
random magnitude which was exerted at a distance 
from the center of the platen proportional to the error 
in heading. 

The contractors built a flight test table which was 

































ORGANIC TARGET SEEKING 


201 


controlled so as to point at a moving target under the 
guidance of a conditioned bird. Hunting with this 
flight test table was not unduly serious, the bird being- 
able without great difficulty to keep the target within 
2 or 3 degrees of the center of the screen. There was 
considerable doubt, however, as to whether the pa¬ 
rameters of the flight test table closely matched those 
of a glide bomb in flight. 

The program was given up because the mechanical 
engineering problem of developing an appropriate 
servo link seemed too difficult in view of the available 
personnel for solving such problems and the high 
promise of success with Pelican and Bat. The Divi¬ 
sion is not prepared to recommend whether further 
study along these lines should be made. Certainly, 
however, should the suggestion be made in the future, 
it should be examined on its merit. The program of 
conditioning organisms is far from a difficult one to 


solve. The problem of devising a stable nonoscil- 
latory servo system is difficult. Whether it is more 
difficult to solve than some of the servo systems de¬ 
veloped during this war for fire control and for other 
purposes is debatable. The frequency of the response 
of the bird was dependably three per second. This is 
admittedly marginal for a rapidly moving missile. 

The experience of the Division, so far as it is con¬ 
clusive, would point to the general observation that 
an organic system of control should not be rejected 
simply because it is organic. Investigators in the 
physical sciences are inclined to discount unduly the 
findings of their colleagues in the field of psycholog¬ 
ical behavior. Such an attitude is far from scientific, 
and there is implicit in the success of the Japanese 
program with organic homing systems the suggestion 
that further study in this field might well be profit¬ 
able. 



Chapter 10 

SIMULATION AND TRAINING AIDS 


101 INTRODUCTION 

F rancis Bacon is quoted as having cited as the 
greatest need of mankind “a machine to aid the 
mind in thinking as a tool aids the hand in working.” 
Mathematics is such a machine. Without its aid 
quantitative thought becomes almost impossible; in 
the field of science and technology, general thinking 
not reduced to quantitative evaluation is at a very 
low level indeed. In many cases, however, ordinary 
mathematics breaks down in the solution of problems 
in research and development. In the field of manual 
activity, hand tools become inadequate where the 
scope of effort to be expended exceeds what can eco¬ 
nomically be supplied by the energy of workers. 
Where hand tools are inadequate, machine tools sup¬ 
plement or replace the energy available from the 
operatives. 

In research and development, problems similarly 
arise where the scope of mathematics, as ordinarily 
conceived, is inadequate for the solution of the prob¬ 
lem. Sometimes this inadequacy finds its expression 
in a loss of economy. Thus, when many repetitive oper¬ 
ations are to be performed, the computing and tabu¬ 
lating machines of Hollerith and his successors per¬ 
form analytical processes more rapidly and with more 
accuracy than they could be accomplished through 
the effort of many workers. 

Problems in mechanics find their most succinct 
statements in the differential equation. A numerical 
solution, however, is possible only if the function in¬ 
volved in the differential equation can be integrated. 
The solution can be obtained only with considerable 
difficulty if the coefficients of the derivatives are not 
linear. These are various limitations in the applica¬ 
tion of mathematics to engineering problems, so that 
the ruse is frequently adopted of considering the rela¬ 
tionship between the elements of a problem as being 
linear in the region in which a solution is sought. Such 
simplification, necessary though it is to permit appli¬ 
cation of conventional mathematics, is frequently un¬ 
justifiable. The naval architect’s curve expressing the 
relationship between driving horsepower and the ra¬ 
tio of the square root of the length to the speed of a 
hull is a well-known example. This curve commonly 
has a sharp break in the region adjacent to the value 


of unity for the speed-length ratio, and it is precisely 
in this region that the interest of the naval architect 
is most acute. 

The product integraph 1 and its successor, the dif¬ 
ferential analyzer, 2 were developed at MIT to meet 
exactly this need. With these powerful tools differ¬ 
ential equations can be solved, if their order is not 
excessive, provided only that the coefficients of the 
derivatives in the equation can be presented graphi¬ 
cally. Thus, not only is the limitation of linearity 
avoided, but a further extension of power is made in 
that it is unnecessary to express the derivative co¬ 
efficient in algebraic language. Empirical data from 
the wind tunnel or from the model towing basin can 
be used directly without recourse to an approximate 
algebraic fit of the experimental data. 

These aids to analysis are simulative in nature. 
They consist of electrical and mechanical circuits 
which can be so adjusted that their performance ex¬ 
actly solves the differential equation under consid¬ 
eration. They have been classified as replica com¬ 
puters in contrast with counting computers of the 
Hollerith type. 

Other types of simulative computers are also well 
known. C. A. Nickle 3 constructed in 1924 a transient 
analyzer which explored the transient behavior of 
electrical and mechanical systems by recording oscil- 
lographically the performance of a dynamically sim¬ 
ilar electrical system. In general, Nickle’s analyzer 
was limited to linear systems, although he recognized 
that the introduction into the simulative circuit of 
electronic elements gave a possibility of exploring 
systems with nonlinear characteristics. 


102 GENERAL 

In the analysis of guided-missile performance, sim¬ 
ulation has been a powerful tool of the Division’s 
contractors. Its aid has been invoked to accomplish 
three aims. As an analytical tool, simulators have 
been built by the Division’s contractors to solve per¬ 
formance equations of missile systems and to adjust 
the parameters of design so that the probability of 
successful flight test with the missile would be at a 
maximum. This is not to say that simulative studies 


202 


GULF PHOTOELECTRIC BOMB 


203 


can replace flight tests; rather, they serve as a val¬ 
uable preliminary supplement to flight tests so that 
valuable effort and irreplaceable time are not ex¬ 
pended unnecessarily. This is one of the most valu¬ 
able functions which simulative analysis can per¬ 
form. In general, a simulative test can be performed 
in minutes or, at most, hours, whereas flight tests 
with the missile occupy days and, if the time of prep¬ 
aration and analysis of the test is included, weeks. 

A second function of simulative analysis by the 
Division was the determination of the optimum com¬ 
bat use of the missile and the exploration of tactics 
which could be employed by the bombardment air¬ 
craft using them. The economy of such a method 
where the dynamic similitude is authentic is obvious. 

The third application of simulation by the Division 
was in the development of training devices intended 
to go in the field with the combat crew, continually 
to supplement their training and combat experience. 
Such a device could also perform an extremely useful 
function in basic training by giving the bombardier 
his first experience with the peculiarities of the new 
weapon. 

Each of these three functions permits a different 
standard of construction. For the analytical device, 
to be used only by trained laboratory personnel, de¬ 
vices may be crude, yet extremely valuable. Acces¬ 
sibility of components and adaptability to variations 
of parameters add to the power of the simulator in 
such cases and are more readily achieved when the 
design is fluid than when it is frozen. To investigate 
possible tactics, the engineering must be sufficiently 
refined to convince the operating personnel of the 
authenticity of the simulation. The construction 
must be good enough to allow the average technician 
to keep it in operating condition. Lastly, the training 
device must be light, portable, and well engineered 
to operate correctly and reliably under adverse con¬ 
ditions. 


103 GULF PHOTOELECTRIC BOMB 4 

In the experiments leading to Azon, Razon, and 
Felix, the question arose as to the practicability of 
any homing mechanism; whether a target seeker at¬ 
tached to a wingless bomb could ever supply suffi¬ 
cient stable control, utilizing the limited available 
amount of lift, to bring the bomb to the target. As 
pointed out in Section 9.2.1, the answer to the gen¬ 
eral problem was sought by using the simplest, most 


foolproof target seeker available, one utilizing a pho¬ 
toelectric cell. 

The general principle of operation of the Gulf 
bomb, that the bomb body supplies most of the lift, 
left some problems to be settled in the design of the 
mechanism. Preliminary experiments with the test 
table, for example, showed that oscillations of dan¬ 
gerous amplitude could be set up if the target seeker 
were rigidly aligned with the axis of the bomb and 
on-off control used. This effect was due to the spring¬ 
like reaction between rudder deflection and yaw an¬ 
gle. After a sudden reversal of rudder setting, it took 
some time for the bomb body to assume a trim angle 
of attack. During this time it was oscillating, and the 
target seeker was therefore yielding alternating error 
signals. Later investigations dealt with the type of 
control that the preliminary analysis had shown to 
be most suitable—the target seeker directed by 
“ears” to point in the instantaneous direction of 
travel, and with proportional control. As the coupling 
ratio of ears to rudders, range of proportional con¬ 
trol, and spring and damping constants were varied, 
the stability of the bomb in yaw and pitch was ob¬ 
served. 


10 31 The Test Table 

This table (Figure 1) was built for the purpose of 
simulating angular motions of target-seeking bombs 
in yaw and course. One of the photoelectric target¬ 
seeking units identical with those in the six bombs 
was mounted on the top platform, which rotated in 
accordance with the motion of the bomb axis. The 
lens was coupled as shown (lens coupling linkage) 
directly with the large spur gear. The latter, driven 
by the course motor, rotated according to the varia¬ 
tion of the bomb in course, right or left. The three 
light beam projectors shown in the photograph were 
used to show, respectively, rudder motion, yaw mo¬ 
tion, and course motion of the bomb. 

A few liberties had to be taken to reduce the num¬ 
ber of factors to be simulated. The parameters nec¬ 
essary to account for the effect of gravity and for va¬ 
riation of aerodynamic behavior with bomb velocity 
were maintained constant. The yaw displacement was 
spring-coupled with the rudder deflection to simulate 
the aerodynamic coupling between rudder and yaw 
in flight. Similarly, as shown, the yaw motion was 
subjected to viscous damping by means of a paddle 
wheel moving in oil. 



204 


SIMULATION AND TRAINING AIDS 



Figure 1 . Gulf target-seeking test table, 












THE GOLD BUG 


205 


In the vector diagram shown in Figure 2, 
\f/ = course error angle; 
a = angle of attack; 

<j> = angle of bomb above line of sight; 
e = target bearing; 

then 

€ = <f> — a — i/'. 


The equations the simulator solves are derived as 
follows: 

la + na + k\a = k 2 8 (1) 

In the above 

I = moment of inertia 
H = aerodynamic damping coefficient 
ki = spring coupling coefficient representing the 
stabilizing moment 
8 = rudder angle 

k \a trim 


k 2 = 


or 


8 = k s e when —5 ^ e ^ +5 
8 = — 20 degrees when e < — 5 


( 2 ) 


and 

8 = 20 degrees when e > 5 

and a — <f> = k A a, 

when k — ^ er cen ^ coupling — 100 

100 


(3) 


By substituting in equation (1), within the range of 
proportional control, 


la + na + kia = k 2 k$e (4 ) 

or beyond that: 

la + fxa + k\a = k 2 - 20 (4a) 

= k 5 a (5) 


Equations (4) and (4a) were solved by the simulator. 
Equation (3) was set up by adjusting the coupling 
between the lens and the course gear, n, ki and ki/k 2 
were varied by adjusting the stiffness of the spring 
and the viscosity of the damping oil. 


10 3 2 Results 

The yaw stability was found to be sensitive to the 
value of the percentage of coupling between the lens 
and the course. (100 per cent coupling is defined as 
the case in which the motion of the lens exactly com¬ 
pensates for the change in the field of view caused by 
yaw motion.) Very small decrease below 100 per cent 
coupling was found to result in yaw oscillation, 


whereas overcoupling gave stability, but at the pen¬ 
alty of decreasing the range of error angles for which 
the rudder displacement can be kept proportional to 
the target displacement. In turn, this proportionality 
was necessary to prevent hunting (see Section 9.2.2). 
The curves in Figures 3 and 4 show the effect of vary¬ 
ing the coupling ratio. 



Figure 2. Vector diagram of homing bomb. 


As it affected the design of the bomb, this observa¬ 
tion meant that it was essential for stability that the 
lens (which was directed by ears projecting into the 
wind stream) move exactly the amount to compen¬ 
sate for pitch and yaw motion of the bomb. This was 
accomplished by adjustment of the linkage between 
the ears and the lens. 

This test table had its faults: it did not yield its 
own record except as the observer watched the lights 
move, its damping was partly frictional and only 
partly proportional to velocity, and there were indi¬ 
cations that the coupling between the lens and the 
course indicator was not free from backlash. But with 
all its approximations, it gave evidence which was 
used in adjusting the actual bomb for field tests. In 
turn, five out of six bombs so tuned made good scores 
on the target, a record unique in the Division's ex¬ 
perience. As a by-product, this table served as the 
basis for the design of the Felix simulator. 

104 THE GOLD BUG 5 

Early work on the test table described above 
showed the general feasibility of the method of con¬ 
trol adopted for Felix, the MIT heat-homing bomb. 
Because of the bulk of the apparatus, the direction 
of view of Felix could not be shifted readily by means 
of ears; instead, the scanning element was coupled 
mechanically with the rudders and elevators. By 
using the known value of the trim angle of attack 








206 


SIMULATION AND TRAINING AIDS 



0 1 2 3 4 5 6 7 8 9 10 II 12 13 14 15 16 17 18 19 20 21 22 

TIME IN SECONDS 


Figure 3. Response of test table simulating target-seeking bomb; 100 per cent ear-to-lens coupling, +5-degree range 
of proportional response. 


produced by a given rudder setting, and hence the 
direction of flight with respect to the bomb axis, it 
was possible to couple the rudders and target-seeking 
eye in such a manner that the latter was always 
directed along the line of motion at equilibrium. To 
the extent that equilibrium was reached without os¬ 
cillation this represented the ideal case, in which it 
was learned that the method was basically sound. To 
insert the oscillation about the trim position, com¬ 
putation was resorted to. 

Eglin Field tests of Felix showed that the general 


principles seemed sound but that details needed cor¬ 
rection. It was realized that manual computation of 
the effects of altering each of the variables was im¬ 
possible, and MIT built a simulator using the Gulf 
model design as a basis but capable of giving more 
information. The main improvement was the addi¬ 
tion of recording pencils to draw curves showing the 
oscillations of the various vectors. 

With the same terminology as in Section 10.3.1, 
the equations to which the Gold Bug conformed 
follow. 



o 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 

TIME IN SECONDS 


Figure 4. Response of test table simulating target-seeking bomb; cases illustrating under- and over-coupling. 



















































































































































THE PELICAN SIMULATOR 


207 


Coupling: 

<f> = k F 8 ( 6 ) 

k F = coefficient of coupling (controlled) 

Bomb body: 

/(7 + $ a) + Ra + k(a — 5) = 0 (7) 

7 = angle between vertical and line to target 

Velocity vector: 

7 + ^ = k L a (8) 

k L = lift factor 

Rudder: 

8 = zigzag, with reversals shortly after target 

crosses line of sight 

Angle of attack: 

la + Ra + ka = k8 (9) 

R = Ik L + R' 

k = spring coupling coefficient 
R' = aerodynamic damping 


It will be noted from the schematic diagram (Fig¬ 
ure 5) that gravity is inserted as a factor, and that 
seven variables are recorded. 



1041 Construction 

The photograph (Figure 6) shows the construction. 
The recording drum and attached pencils with their 
piano-wire drives are easily recognized. Heavy cylin¬ 
drical weights supply the inertia. The scanner is near 
the top center of the picture, and the target is under 
the wooden housing to the right. 

Figure 7, taken from the opposite side, pictures the 
inertia weights, the coupling springs, the five arms 
which drive the recording pencils, and the differential 
which operates the course arm'. 

Reference to Figure 7 reveals how little engineer¬ 
ing refinement is required to build a simulative mech¬ 
anism of great analytical power. In spite of its crude 
appearance, it was thoroughly adequate to give quan¬ 
titative answers which determined the optimum 
coupling ratio between the control surfaces and the 
scanning system, and to evaluate the permissible 
overall time lag within the precision with which the 
fundamental characteristics of the missile could be 
determined. 

10 4 2 Results 

Representative curves as drawn by this device are 
shown in Figure 8. It will be seen that as the target 
was moved manually, the elevators and bomb body 
oscillated about an average position, following the 
target motion, and the velocity vector swung around 
smoothly. 

It is safe to say that not until this machine was put 
into operation was it possible to settle design factors 
upon a basis better than an inspired guess. 

10.5 THE PELICAN SIMULATOR 6 

In the two simulators described above, the com¬ 
putations were carried out mechanically. The investi¬ 
gators on the Pelican project found it profitable to 
combine electrical and mechanical computation in 
their study of longitudinal (pitch) stability. The con¬ 
trolling equations have been derived in Section 1.4. 

1051 Construction 

The circular table (Figure 9) was free to rotate 
about a vertical axis. Angular motion was damped 
by a direct-connected oil-filled cylinder with close 
clearances, so that the damping was directly propor¬ 
tional to angular velocity. 






















208 


SIMULATION AND TRAINING AIDS 


Geared to the table were two synchros. a The first, 
the table synchro, produced a voltage proportional 
to the angular displacement of the table from the 
neutral position. The second, the torque synchro, ap¬ 
plied a driving torque of a magnitude and direction 
determined by the gyro mounted on the table, the 
servo link under test, and the computing circuit (Fig¬ 
ure 10). The third synchro shown in this figure was 
mounted on the output of the servo link so that its 
voltage was proportional to servo output velocity. 
The torque applied to the table then becomes: 

L = KE g (10) 

where E 0 is the voltage between the two grids of V3. 
Now if 


a Synchro is a word coined by the Navy to indicate a 
dynamo-electric machine having a single-phase rotor and 
three-phase stator. Specifically, Bendix Co. Autosyns were 
used. 


I = moment of inertia of the table system 
fx = damping coefficient of the table system, 


then the motion of the table will be exactly identical 
to the pitching action of the glide bomb when con¬ 
trolled by the same gyro and servo link, provided the 
following relationships between the glide-bomb char¬ 
acteristics and the circuit parameters are maintained. 

I j = — Mq — Mat 


R£i R»C, a 

III K j- - Mac + aMct 

KC 

IV = b(- Ma + aMa]) 
V y = Mg - bMa 

VI y 4 = Mj 


(ID 



Figure 6. Gold Bug calculating machine, front view. 











THE PELICAN SIMULATOR 


209 


In the above, M = angular acceleration in pitch 
dM dM dM 

da d8 dq 

dM dM 


^ is the velocity head = 
a is the angle of attack 
5 is the elevon angle 

A procedure for adjusting the circuit to satisfy 
equations (11) is given in the investigator’s report. 


10 5 2 Operation 

The entire assembly as it was used in studying the 
performance of the gliders constructed by this con¬ 
tractor included (Figure 11): the test table (center), 
a servo unit (left) in which torsion rods simulated the 
aerodynamic loads on the elevons, and the power 


unit (right). Performance was recorded by a record¬ 
ing oscillograph (Brush Oscillograph, Model OBC). 
Oscillations in <j> , a, and b were recorded for different 
settings of the minimum angle 4> which saturated the 
differential amplifier, the gyro rate dL/da, and dL/db. 

10 5 3 Results 

In this device, simplicity of construction and neat 
appearance were achieved by converting the calcu¬ 
lated outputs to voltages proportional to the angles 
to be measured. The advantages are apparent; the 
longer time required for calibration was more than 
compensated for by the greater stability once adjust¬ 
ments were made. 

With this, as with the others already described, it 
was determined that slight variation of some of the 
design factors (<£, gyro rate, L a , and Lt) from the 
adopted values could produce unsatisfactory oscil¬ 
lation. Here again the time scale was tremendously 



Figure 7. Gold Bug calculating machine, rear view. 










210 


SIMULATION AND TRAINING AIDS 



shortened by the transfer of hundreds of “flights” 
from the test field to the laboratory table. 


106 PELICAN SERVO TEST TABLE 7 

Chapter 7 included a discussion of the servo sys¬ 
tem developed at MIT for use in the Pelican series 
of glide bombs. The stages in the work of this con¬ 
tractor were: 

1. Mathematical analysis of a proposed system and 
determination of the required dynamic and static 
characteristics of each component; 

2. Study of the system on a flight table designed 
to simulate the more important dynamic and static 
characteristics of the glider and radar system; 

3. Design of components; 

4. Dynamic and static checks of each component; 


5. Rechecks upon the flight table of the control 
system comprising the components finally adopted; 

6. Flight tests of the complete missile. 

Items 2 and 5 entailed the use of a mechanical 
flight table that actually simulated the Pelican glider 
but was capable of adjustment for other missiles of 
similar characteristics. It should be mentioned that 
this study was the only one in which the entire closed 
loop was explored as a system, including the servo¬ 
mechanism, airframe, and target seeker. 

The flight table as built by this contractor was de¬ 
signed to simulate the missile in roll and in pitch. 
Figure 12 is a functional diagram of the complete 
test table with its associated equipment. It will be 
noted that this bore some similarity to the table de¬ 
scribed in the preceding section but allowed for the 
introduction of two additional variables—roll orien¬ 
tation and action of the radar receiver. Rotation of 
























































ELECTRONIC SIMULATORS: AZON AND RAZON 


211 


the table top about horizontal and vertical axes simu¬ 
lated roll and yaw motion of the glider, respectively. 
Up-and-down motion of the target provided for the 
equivalent of pitch motion. 

Portions of the mechanism which were not sensi¬ 
tive to the motion of the table, such as the elevons 
and their servomotors, were located conveniently 
near-by. 

The equation of motion of the glider in roll was: 

Idbd — L p <p = A 4> (12) 

dd = differential displacement of the elevons; 
l d = roll torque for unit elevon displacement; 
L p = damping torque for unit rate of roll; 

</> = roll angle; 

A = moment of inertia about the roll axis. 

An exact simulation of equation (12) was obtained 
by designing the table so that: 

A/R = moment of inertia of the table plus its mo¬ 
tor and gearing referred to the table; 

Lp/R = effective damping of the table; 

l d /R = torque produced upon the table by unit 
displacement of the simulated elevons; 

<t> = roll angle of the table; 

R = a scale factor. 

A variable-displacement hydraulic pump whose 
displacement was a function of the roll angle of the 
table provided a rate of yaw proportional to the an¬ 
gle of roll. It will be remembered from Chapter 1 that 
the Bureau of Standards glide bombs were unique 
among the missiles of the Division in that they turned 
in yaw only as the result of roll. 

The contractor's report 7 describes how the appara¬ 
tus was arranged to make the above relationships 
valid. With the flight table, most of the conditions 
met in flight were adequately represented. Important 
measurements made with the use of this table were: 

(1) accuracy with which the glider holds its course; 

(2) damping of the glider to suddenly applied roll 
moments; (3) the nature of the response to disturb¬ 
ances in flight direction. 

Two general systems were studied, one using rate 
gyros only for stabilization, and the other using free 
gyros and rate information obtained electrically. The 
results eliminated the first type in favor of the second. 

1061 Results 

Engineering selection of components was based 
upon their behavior on the test table. In addition, 


the entire servo system was tested thoroughly before 
and after installation in the glider. The effectiveness 
of this procedure was reflected in the performance, as 
described in Chapter 7. 



Figure 9. Pelican pitch simulator. 

107 ELECTRONIC SIMULATORS: 

AZON AND RAZON 

Philbrick has discussed 5 the fundamental princi¬ 
ples of electronic or mechanical-electronic simulation, 
and related the history of the collaboration between 
Division 5 and Division 7 on this subject. It suffices 
to say here that, as the result of work upon bread¬ 
board models at Columbia University 8 under Section 
7.2, it was shown that the differential equations for 
the flight of Azon, Razon, and Roc were amenable 
to approximate solutions by mechanical-electrical 
means. Some work at Columbia also showed the 
feasibility of stable control of the television Roc and 
the necessity for proportional control for Roc. Re¬ 
duction of breadboard models to useful training in- 


b Aiming Controls in Aerial Ordnance, G. A. Philbrick 
Division 7, Volume 3. 




TO TABLE SYNCHRO TO SERVO SYNCHRO 


212 


SIMULATION AND TRAINING AIDS 




Figure 11. Test table assembly. 


to rower supply 




























































































































ELECTRONIC SIMULATORS: AZON AND RAZON 


213 



struments was accomplished by Division 5 con¬ 
tractors. 


1071 Mathematics 

The equations representing the motion of Razon 
may be expressed (in rectangular coordinates) as 
follows: 


- jg %Cl(s,)v* cos e 

(13) 

- ~ §C,(i)V 

(14) 

range coordinate; 



y = azimuth coordinate; 

A = area of cross section of bomb; 

M = mass of bomb; 
p = air density at altitude h; 

V = velocity of bomb tangent to trajectory; 

6 = angle between vertical and tangent to 
trajectory; 

Cz,(5 e ) = ballistic coefficient as function of elevator 
angle; 

Cs(br) = ballistic coefficient as function of rudder 
angle. 

We may simplify these by expressing pV 2 cos 0 and 
pV 2 as average functions of time, and by neglecting 
cos 0 in order to permit use of the same linkage for 


both range and azimuth channels. The above equa¬ 
tions then reduce to: 


x=f(t)8 e (15) 

y=f(t)8r (16) 

All the constants in equations (13) and (14) have 
been absorbed in the time function, which is very 
nearly a square law, i= Kt 2 . Adding the per¬ 
spective effect, the bomb deflection as seen by the 
bombardier is, if z = the altitude of drop, 

x x 


Ox = tan 


—i 


- h 

y 


Oy = tan 1 —-—7 = - 
y z — h z 


z — h 

y 


(17) 


- h 


10 7 2 The Computer 

For Azon and Razon a simple on-off control stick 
was used. Application of control caused the bomb 
rudder to move at a constant rate of 33 degrees per 
second to a limiting value of 20 degrees. In the com¬ 
puter this was simulated by applying a voltage of 
proper polarity to an integrating amplifier and a 
diode limiter. The output of this amplifier therefore 
yielded 8 e or 8 r , according to whether range or azi¬ 
muth was under consideration. 

Two different methods were used in inserting the 
time function. In one, the potentiometers concerned 




















































214 


SIMULATION AND TRAINING AIDS 


were wound on 10-to-l taper cards, which approxi¬ 
mated the square law reasonably well. Only one man¬ 
ufacturer, however, proved capable of winding po¬ 
tentiometers of the high resistance and steep taper 
required, and he was unable to supply adequate 
numbers. For all later models, linear potentiometers 
were used, driven by the simple four-bar linkage (Fig¬ 
ure 13) described by Philbrick and so designed as to 
rotate the potentiometer shaft at a rate proportional 
to the square of the rate of turn of the synchronous 



Figure 13. Square-law four-bar linkage. 


motor driving it (Figure 14). The result is to multi¬ 
ply 8 by Kt 2 . 

x was obtained from x by double integration in in¬ 
tegrating amplifiers. The output from amplifier IV 
was a voltage proportional to the azimuthal displace¬ 
ment of the guided bomb from its unguided trajec¬ 
tory. Lastly, the variable scale factor, which was pro¬ 
portional to the distance between the bomb and the 
airplane, was introduced. This was also proportional 
to /(£), so another potentiometer on the same shaft 
allowed multiplication of the gain of the last ampli¬ 
fier by the factor l/(z — h) = l/[f(t)]. 9 

10 7 3 Model 1010 

The output was presented in one of two ways. In 
the Model 1010 trainer, the output voltage for each 
coordinate was fed to one pair of field coils of a spe¬ 
cial galvanometer unit. This unit was a Mark 18 
gunsight, modified by the replacement of the stand¬ 
ard trail coils with special high-resistance coils. In 
this unit, a gyroscopically mounted mirror was pro¬ 


cessed by a voltage applied to the trail coils, the 
effect being that of a galvanometer capable of motion 
in two coordinates. 

A spot of light reflected by the galvanometer mir¬ 
ror appeared superposed on the target picture pro¬ 
vided by an Arn^ A-5 or A-6 bombing trainer (Fig¬ 
ure 15). In effect, the bombardier, looking through 
the Crab sight, saw the bomb flare superposed on the 
target. Control by the standard stick resulted in ap¬ 
propriate scaled motion of the bomb. Suitable relays 
permitted automatic stopping of computation (for 
scoring purposes) at the end of the time of fall, fol¬ 
lowed by recycling to prepare for another run. 

The flare projector, seen enclosed in the vertical 
rectangular box to the operator’s right in Figure 15, 
is shown in detail in Figure 16. The computer box is 
shown in Figures 17 and 18. 

The A-5 and A-6 trainers simulated the usual pro¬ 
cedure of the bombardier in standard bombing, with 
a target photograph projected upon a white surface 
for observation through the bombsight. The equip- 



0 60 120 180 240 300 360 

DEGREES ROTATION OF POTENTIOMETER SHAFT-► 

Figure 14. Approximation of square law by linkage. 


ment for operating the bombsight was standard, and 
the response of the field of view to the bombardier’s 
controls reproduced the actual target as it would be 
seen from a moving plane. To superpose the Razon 
attachment required addition of the flare projector, 
the control stick, the computer box, and the Crab 
sight. 

Adjustments needed were alignment of the flare 
projector with the impact light on the trainer, and 
balancing of the ten amplifiers in the computer box. 



































ELECTRONIC SIMULATORS: AZON AND RAZON 


215 



Figure 15. Model 1010 trainer in operation 



216 


SIMULATION AND TRAINING AIDS 










I,—. 





= 

OPTICAL AXIS W 


h" & 

;8 NEGATIVE LENS 















1 - 

4 S 

J 

"Tr 


26" F.L. PROJECTION LENS 


40" 


Figure 16. Optical arrangement of flare projector. 




(Directions for the latter are seen engraved on the 
face of the computer panel.) 



Figure 17. Computer box for Model 1010 and Model 
1020. 


Pushing the “reset” button allowed the motor seen 
at the bottom center in Figure 16 to drive the linkage 
to the “zero time” position. From then on, all the 
operations of the bombardier were identical with 
those he would perform in combat: setting and syn¬ 
chronization of the Norden bombsight, automatic 



Figure 18. Interior of computer box. 

bomb release, and use of control stick to keep flare 
superposed on the target image as seen in the Crab 
mirror. At impact the flare light remained fixed, in¬ 
dicating where the controlled bomb would have hit, 
and the impact light showed where the bomb would 






























































































ELECTRONIC SIMULATORS: AZON AND RAZON 


217 


have hit if uncontrolled. With expert bombardiers 
the latter point was so near the target that it was fre¬ 
quently desirable to direct the bombardier to syn¬ 
chronize on one target and then steer the bomb to 
another target. 

The dynamics of the moving spot suggested by the 
trainer in response to control-stick manipulation are 
very close to those of the apparent motion of the 
Razon flare as seen from the bombardier’s position. 
Except, therefore, for the discomforts of flight, 
cramped position, noise, cold, altitude, etc., the 
Model XI010 is a very realistic trainer. A very short 
program with this device is equivalent to a good 
many flying hours. Installed first at Fort Dix and 
later at the Columbia (S.C.) Air Base, it was used in 
training Azon and Razon crews. (For Azon, of course, 
the range section of the device was not used.) A simi¬ 
lar trainer has been made available to the Navy and 
is under analysis by them at Traverse City, Michigan. 


10 7 4 Oscilloscope—Model 1020 

A portable model was also devised for briefing 
bombardiers with actual target photographs. In this 
model (Figure 19) the output voltages from the com¬ 


puter were fed to the deflection plates of a cathode- 
ray oscilloscope. The target photograph, mounted in 
the illuminated box seen projecting from the left 
front of the scope, was rendered coplanar with the 
face of the scope by the 45-degree semireflecting glass 
mounted on the front of the scope. The operator 
pushed the release button on the front of the com¬ 
puter and maneuvered the flare spot to the chosen 
target. The spot positioner allowed the instructor to 
insert any desired aiming error. 

This device, while less realistic than the Model 
1010, was portable and required no special photo¬ 
graph. Any photograph or transparency printed to 
the proper scale (1 in. = 1,000 ft) could be utilized. 

10 7 5 Results 

With early versions of this equipment it was de¬ 
termined that: (1) Razon with on-off control was 
stable; (2) Roc 00-1000-V required proportional con¬ 
trol for guidance; (3) bomb-guiding techniques could 
be acquired within a relatively short training period; 
(4) the fact that the design involved several approxi¬ 
mations militated against its use as a quantitative 
device, as had been proposed, but did not invalidate 
the above conclusions. 



Figure 19. Model 1020 trainer in use. 



218 


SIMULATION AND TRAINING AIDS 


10 8 MIMO ELECTRONIC SIMULATOR 10 

Remote control of Roc, using television informa¬ 
tion, was the subject of extended investigation. Pre¬ 
liminary work under Section 7.2, reported by Phil- 
brick, had shown that an unmodified pursuit course 



Figure 20. Geometry in vertical bomb-target plane. 



Figure 21. Circuit for addition and subtraction. 


must result in a miss of a moving target, but that a 
“regulator” for converting the pursuit to a collision 
course would give good results. The Columbia simu¬ 
lator 8 also showed that (1) satisfactory steering was 
accomplished only after considerable practice, and 


(2) both experienced and inexperienced operators 
improved their scores by continued practice. 

By the time television equipment for Roc was 
available, the contractor’s representatives had ac¬ 
quired the necessary skill with the aid of their test 
cart (Section 4.6.2), but it was manifestly impractical 


R 2 



c 



to waste both bombs and time training Army bom¬ 
bardiers with the actual materiel. The Division, 
therefore, arranged for the building of a trainer fol¬ 
lowing the general outline of the Model 1020 Razon 
trainer, modified to conform to the Roc system of 
control (proportional) and to present the results from 
the point of view of the nose of the missile. 

















































MIMO ELECTRONIC SIMULATOR 


219 


General Principles 


In Figure 20, 

r = angle from vertical to line of sight from bomb 
to target; 

7 = inclination of bomb axis from vertical; 

X = angle between line of sight to the target and 
the axis of the bomb; 
r = slant range to the target; 

V p = bomb velocity in air mass; 

V e = target velocity in air mass; 

L = aerodynamic lift; 

M = mass of bomb; 
g = acceleration of gravity; 

K = control surface constant; 

8 = control surface angle; 
p = air density; 

S = area of control surface. 

Then for use in the simulator: 


r = 7 + X 
7 T = Vp\ + V e 


. = L _ 91 

7 MV P V p 


(18) 

(19) 

( 20 ) 


The above equations are linear, derived from the 
actual equations 11 by introduction of the following 
simplifying approximations. 


1 . (a) sin t = r 

(b) cos r = 1 

(c) sin 7 =7 


(Equation 19) 
(Equation 19) 
(Equation 20) 



1 SVp'Cl 

2 p MV P 


= KlpV p 6 


r 2 



3. Vp, r, and p are linear functions of time after 
release and are not affected by actual control excur¬ 
sions about the assumed course. 

Of the above approximations, 1 (a) is valid, since 
X is always small; 1 (b) seems unjustified, but V e is so 


b b' 



1 FOR BALANCED OUTPUT, USE b AND b‘ 

2 FOR BALANCED INPUT, USE d AND d’ 

3 FOR SINGLE-ENDED OUTPUT, CONNECT b OR b‘ 
TO GROUND 

4 FOR SINGLE-ENDED INPUT, CONNECT d OR d 
TO GROUND 

5 FOR AMPLIFICATION WITHOUT CHANGE IN SIGN 
(•(-A) USE d AND b, OR d'AND b' 

6 FOR AMPLIFICATION WITH CHANGE IN SIGN 
(“A) USE d AND b' ( OR d'AND b 

Figure 25. Basic amplifier circuit. 












































220 


SIMULATION AND TRAINING AIDS 


—-TIME 



much smaller than V p that substitution of V e for 
V e cos r has only a slight effect upon r; 1(c) is sub¬ 
ject to similar reasoning— gy/V p and g sin y/V p are 
both small numbers with little effect upon the first 
term of the equation. 

For simulation, a voltage corresponding to X was 
applied to the deflection plates of an oscilloscope. 

10 8 2 Electronic Computation 

Electronic solution of the above equations and ap¬ 
plication to the design of the simulator were based 
upon a few general principles which are not new but 
are not too well known, so they will be discussed 
briefly. 


Addition and Subtraction 

The circuit in Figure 21 may be used either for 
addition or subtraction, depending upon the polarity 
of the applied voltages. The block in each of these 
circuits represents a d-c amplifier, with gain /x. 

The exact input-output relationship is 

— (Eg -f- Eb -f~ E c -f- • • •) 

i+- 


If ix is high enough, this reduces to a simple alge¬ 
braic summation. 


10 8 4 Multiplication and Division 

Similarly a single circuit (Figure 22) can be used 
for multiplication and division. In this case 

R2 En 


E 0 = - 


*P+7( 1 + f)] 


( 22 ) 


Again, if n is large, this reduces to the approximation 


E„ = | 


(23) 


10,8,5 Time Integration 

The circuit for integration (Figure 23) is better 
known. For this arrangement, 

Eodt = CR 1 (l'+ii)f E ‘ dt ~CR l (l + n)l Eodt (24) 

Again, if /x is high enough, and, better, if E 0 re¬ 
verses occasionally or periodically, this becomes very 
nearly 



E 0 = 


(25) 




























































MIMO ELECTRONIC SIMULATOR 


221 


Time Differentiation 


Interchanging R i and C gives the usual circuit for 
differentiation (Figure 24). 

Here the exact statement is 


E o = 


nCRi 
1 + M 


E n 


CRi 


1 +M 


- E o 


(26) 


If is large enough, this becomes the desired dif¬ 
ferential 

Eq = — R 2 CE 0 (27) 


8 7 Roc Computer 

The principles deduced above indicate that the 


higher the value of /x, the greater the accuracy of 
computation. The amplifiers used (Figure 25) had 
gains of 170 to 500. 

Figure 26 shows how these principles were utilized. 
Amplifiers 5 and 6 simulated the control servo on the 
bomb. The signal indicating the position of the con¬ 
trol handle, combined with 8, the rudder position, 
determined the speed of the servomotor driving the 
rudder. The input to amplifier (6 therefore represented 
8. Limited and summed algebraically, it was applied 
to amplifier 5. , 

Integration of 8 in amplifier 5 yielded 8. This, mul¬ 
tiplied by KpVp in the motor-driven linear poten¬ 
tiometer A, was combined with the g/V v factor to 




Figure 27. Complete simulator model No. 1 with dual control stick (top), with two single-control sticks (bottom). 









222 


SIMULATION AND TRAINING AIDS 


yield y, solving equation (20). This, integrated in 
amplifier 4, yielded y. 

Similarly, f was obtained from the summation of 
XFp and V e and division by r in amplifier 3—see 
equation (19). 

Integration of r in amplifier 1, followed by subtrac¬ 
tion of y from r in amplifier 2, yielded X for applica¬ 
tion to the oscilloscope plates. 



Figure 28. Block diagram of two-coordinate televi¬ 
sion bomb (Mimo) simulator model No. 1 components. 


A separate circuit allowed recording of the miss on 
meters on the computer panel. (Figure 27) 

The block diagram (Figure 28) summarizes the 
above and shows parallel computation for the two co¬ 
ordinates. 

10 8 8 Results 

This contractor adapted the basic ideas developed 
by the Division 7 contractor, Columbia University, 


to the needs of the Roc program. He combined with 
the basic computer developed at Columbia the prin¬ 
ciples of sound electronics engineering, so that the 
resulting apparatus was stable, reliable, and other¬ 
wise suitable for training purposes. The report also 
gives a good statement of the principles to be fol¬ 
lowed in selection of components for similar com¬ 
puters for other projects. 

The presentation is not too realistic. A spot of light 
on the oscilloscope screen represents the target, its 
motion corresponding to the motion of the target on 
the television screen as the bomb falls. The operator 
gets none of the sensation of target approach pro¬ 
duced by increase in size of the image. Nevertheless, 
a large number of tests run by the contractor have 
shown that training is necessary to insure good con¬ 
trol, and that this simulator could provide such 
training. 

A more serious deficiency in this particular device 
is the fact that the design is based upon the assump¬ 
tion that the steady-state coefficients are valid in the 
transient state. Within the existing body of knowl¬ 
edge this was necessary, but as a result the trainer 
failed to reproduce oscillations which were very no¬ 
ticeable in the actual missile. It was possible to 
introduce oscillations to match the characteristics of 
those observed in any given case, but not to deter¬ 
mine what oscillations would arise under different 
conditions. This would be most desirable in future 
training devices, but is dependent upon nonexistent 
(at present) wind-tunnel studies of the effects of 
transients. 

The Roc designers built (Section 4.6.2) a computer 
which was to be mounted on the face of the scope. 
It was to provide a reticle specifying in two coordi¬ 
nates the desired position of the target in a pro¬ 
grammed collision course. The design was analogous 
to the single-coordinate computer attached to the 
contractor's test cart. 

With the Army preparing to test the Mimo-Roc, 
the simulator and computer have been supplied to 
the test group for use in training the bombardiers 
concerned. 



























Chapter 11 

TRANSITION AND ENGINEERING ACTIVITIES 


hi INTRODUCTION 

I N any developmental program the provision of 
competent laboratory research is not sufficient to 
produce a finished usable device. This is equally as 
true in the case of guided missiles or other instru¬ 
mentalities of warfare as it is with automobiles, air¬ 
craft, or any peaceful product. This need is well rec¬ 
ognized in normal peacetime industrial endeavors. 
Indeed, in some large industrial organizations, the 
research laboratory is supplemented by a general en¬ 
gineering laboratory which is charged with the reduc¬ 
tion to practice of the basic principles evolved by the 
research group. Thus, under ordinary peacetime con¬ 
ditions the research group finds its primary activity 
in the discovery of new scientific principles in the ex¬ 
tension of human understanding. 

The engineering group finds its principal activity 
in the application of the new aspects of human un¬ 
derstanding to human use. Even in wartime when 
research laboratories are more concerned with the de¬ 
velopment of devices than the establishment of new 
laws, this need is still acute. The activities of an en¬ 
gineering group in implementing a research organiza¬ 
tion are manifold. Such a group modifies the basic 
design established by the research laboratory so that 
techniques of proved economy and reliability can be 
applied in production of the new device. It schedules 
operations during the initial production phases so 
that the whole project comes to fruition simulta¬ 
neously. Finally, it is able to make wise compromises 
between the ideal requirements of each component 
of a system so that the integrated design of the whole 
system is at an optimum. It compromises the ideal 
requirements of the integrated system with the re¬ 
quirements of rapid and economic production. 

112 GENERAL 

The NDRC philosophy as regards this phase of 
developmental activity envisaged a central office 
charged with the responsibility for all engineering 
activity of each of the nineteen divisions and the two 
panels. This office was to serve as a consulting engi¬ 
neering organization with the divisions and panels of 
XDRC as clients. The needs of the divisions in some 


cases, however, outstripped the capacity of this gen¬ 
eral office to render engineering service, so that Divi¬ 
sion 5 planned early in 1943 to establish a group of 
engineers in Division headquarters. 

These engineers would be assigned to the several 
projects within the Division. During the research 
phases of the project, they would advise the section 
chiefs as to methods of development which would 
lead to readily producible items and assist in explor¬ 
ing possible sources of procurement for the device 
after it had been developed. During this phase of the 
work their responsibilities were advisory only; that 
of the section chiefs was fundamental. 

The device having been proved in test, these en¬ 
gineers were responsible for directing its development 
for production. They acted as continual and author¬ 
itative liaison between the developmental contractors 
and Division headquarters. They maintained con¬ 
tinuous liaison between the developmental contrac¬ 
tors and the section chiefs under whom the research 
originating the device had been carried out, to the 
end that none of the basic principles evolved during 
the research phase of the program would be violated. 
During this phase of the program their responsibility 
was fundamental; that of the section chiefs was 
advisory. 

These philosophies were in no sense contradictory. 
The group of engineers contemplated for the Di¬ 
vision headquarters was not intended either as a 
duplication of or as a substitute for the Central En¬ 
gineering and Transition Office at NDRC headquar¬ 
ters. Rather, it was the intention that this group, 
through close liaison with the central engineering 
organization and by virtue of its continuing associa¬ 
tion with the projects of the Division, would be able 
to draw from the headquarters office help and advice 
necessary to meet the peculiar requirements of each 
Division problem as it arose. This pattern of opera¬ 
tion closely followed that of normal peacetime de¬ 
velopmental enterprise. 

Pressure of wartime operation thwarted its com¬ 
plete consummation. It was impossible to assemble 
in the Division headquarters a group of engineers of 
sufficient size and individual competence to carry to 
completion each project instituted by the research 
groups. 


223 


224 


TRANSITION AND ENGINEERING ACTIVITIES 


There is no question that the failure to organize 
a sufficient engineering staff to supervise in detail the 
engineering developments of the programs initiated 
in laboratories resulted in loss of economy in the 
Division’s operation. The importance of economy in 
wartime activities is frequently overlooked. It is just 
as vital in research and development programs in 
connection with warfare as it is in similar programs 
of a peacetime nature. Under conditions of war, how¬ 
ever, economy is measured in the saving of time. 
Funds can be made available in very large quantities; 
time is inexorably limited. 

The experience of every industrial enterprise points 
to the value of lodging responsibility for engineering- 
development in a different group of individuals from 
those associated with research. The lessons learned 
from the experience of these organizations is readily 
transferred to wartime activity. The contract made 
by the NDRC Engineering and Transition Office 
with the New England Power Service Company is 
typical of what can be done to make engineering 
talent available for the advanced phases of wartime 
development. Should the Government in the case of 
another war undertake the mobilization of science, 
as was done in World War II, a similar procedure 
should be followed. It should, however, be initiated 
earlier in the program, and its scope should be 
broader. 

It appears that with a large staff of competent 
engineers made available in this manner, the work of 
the entire organization would be benefited. False 
starts along avenues of unprofitable technique would 
be avoided through the guidance of mature engineer¬ 
ing judgment. Research would be strengthened by 
eliminating the necessity for diffusing the efforts of 
research personnel into fields inappropriate to their 
particular talents. 

113 GLIDE BOMBS 

As has already been stated (see Chapter 1) the 
basic research on the Pelican missile was carried for¬ 
ward under two cooperating groups. The aerody¬ 
namics and servomechanism were developed by the 
National Bureau of Standards under Division 5; the 
radar receiver, by the Radiation Laboratory at MIT 
under Division 14. Similarly with Bat, research was 
carried forward on the radar equipment by the Bell 
Telephone Laboratories under contract with the 
Bureau of Ordnance and the Navy, while the air¬ 
frame and servo development was carried forward 


by the National Bureau of Standards and by the 
Servomechanisms Laboratory of MIT under the di¬ 
rection of Division 5. 

The welding of these research programs into a 
workable missile was in each case an engineering- 
project of some magnitude. Under its contract with 
Massachusetts Institute of Technology (OEMsr-240) 
the Division established an engineering group at the 
National Bureau of Standards known as the MIT 
Field Experiment Station. While this group had a 
certain fundamental research responsibility, it was 
recognized by the Division that the major problems 
remaining to be solved were technological rather than 
scientific in nature. Senior responsibility for the ac¬ 
tivities of the group, therefore, was lodged in an 
engineer rather than in a research scientist. A small 
group of associate engineers and a few draftsmen sup¬ 
ported him, and under his direction a pilot produc¬ 
tion line in radar equipment and in servo links was 
established with Navy cooperation supplying per¬ 
sonnel. 

This same group undertook and discharged the 
responsibility for coordinating the design of the air¬ 
frame with that of the servo link and for seeing that 
all items scheduled for production met appropriate 
specifications for performance at an altitude and un¬ 
der conditions of adequate range of temperature and 
humidity. They developed production designs and 
checked them on the pilot assembly line. They acted 
as consultants for the Navy’s supply agencies when 
problems of mass production arose. 

They assisted at early flight tests and supervised 
the tests in the more advanced stages, so that ap¬ 
propriate data obtained in these experimental activ¬ 
ities became embodied promptly in the working de¬ 
sign. All the field test equipment for the initial squad¬ 
ron that used Bat in combat was designed by this 
group and fabricated on the production line under 
its supervision. It provided a competent engineer 
to accompany the Bat missile into the theater, and 
it recalled him to supervise its modification and ex¬ 
tension to the aircraft equipment as suggested by the 
initial combat experience with the weapon. 

114 AZON AND RAZON 

For the engineering- development of the high-angle 
dirigible bomb, the Division made contracts with the 
Union Switch and Signal Company (OEMsr-1081, 
1285, and 1415) under the terms of which the con¬ 
tractor was to take the designs developed by the 


= * 4 


* ir 



FELIX 


225 


Gulf Research and Development Company under 
Contract NDCrc-183 and modify them for produc¬ 
tion and for combat use. The activities of this con¬ 
tractor were directed from the Division headquarters 
office. The Division was supported in this work by 
the consulting engineering service furnished by MIT 
under Contract OEMsr-240 and particularly by the 
advice and assistance of the NDRC Engineering and 
Transition Office. The missile originally designed by 
Gulf to prove the fundamental principle developed 
by MIT only approximated the final geometry and 
loading. It was necessary under this contract for the 
new contractor to devise means of incorporating the 
exact payload required by the Services into the dir¬ 
igible design developed by their predecessor con¬ 
tractors. 

Furthermore, the components used by the research 
group, radio, flare, gyro, and servomotors, had been 
improvised and had not been necessarily intended for 
combat use. Union Switch procured an electric- 
driven gyro to replace the pneumatic-powered one 
used in the early dirigible high-angle bomb program 
by MIT and by Gulf. The original radio receiver (see 
Chapter 6) had been of a type which would produce 
maximum rudder or elevator deflection in case of loss 
of the received signal by the missile. Union Switch, 
through the NDRC Transition Office, procured the 
development of a superregenerative receiver under 
subcontract with the General Instrument Company. 
While this radio receiver was later superseded by a 
more selective model, it had the property of being 
“fail safe,” so that if radio transmission were lost 
during a drop, the error of impact would not be in¬ 
creased by virtue of the radio control feature of the 
missile. 

Power for operating the bomb had been obtained 
from an assembly of batteries—conventional storage 
batteries for the servomotors and for the aileron 
solenoid, dry B batteries for the plate supply of the 
radio receiver. Union Switch procured, through the 
cooperation of the Navy Coordinator of Research and 
Development (now the Office of Research and Inven¬ 
tions), a compact, one-shot, expendable lead and sul¬ 
phuric acid storage battery in a spillproof case (Wil¬ 
lard NT-6). This battery had adequate capacity to 
supply all the power required for the bomb mecha¬ 
nism at the low temperatures likely to be encountered 
on long high-altitude missions. The radio receiver 
was provided with a dynamotor so that the plate 
supply did not require B batteries, whose output 
voltage at low temperatures is very low. 


Early experimental work on the high-angle diri¬ 
gible bomb program had been dogged by unreliability 
of the flare by which the bombardier followed the 
trajectory of the bomb in flight. Under these con¬ 
tracts the contractor investigated other types of py¬ 
rotechnic flares, assisted through the Division head¬ 
quarters office by contractors of Division 11. When 
a satisfactory design emerged, the Army Ordnance 
Department was encouraged ^o undertake its produc¬ 
tion at the Picatinny Arsenal. 

Finally, it was the responsibility of this group to 
work out all the multitude of compromises necessary 
to meet the several, and in some cases conflicting, 
specifications which were deemed to be pertinent. 

FELIX 

The NDRC Engineering and Transition Office was 
the major contributor to the engineering of Felix for 
production. A group of contracts was made by the 
Division to procure engineering development of this 
missile for production. Pressure from the Army Air 
Forces, suggesting an early operational use of the 
weapon, prompted the Division to undertake pro¬ 
duction design before the missile had been completely 
proved in tests. 

Accordingly, one contract (OEMsr-1274) was 
made with the Norton Company of Worcester, 
Massachusetts, for the design and sample production 
of one hundred tails of cruciform structure, although 
it was well realized that roll stability with this type 
of structure posed problems not yet fully solved. A 
succession of contracts with Remington Rand pro¬ 
vided for the production design of tail structures 
having octagonal fins, for nose housings, and for the 
target-seeking assembly. 

Remington Rand subcontracted bolometer supply 
to Cambridge Thermionic Laboratories, Cambridge, 
Mass., and to Eppley Laboratories in Newport, 
Rhode Island. The Heat Research Laboratory at 
MIT had built bolometers of satisfactory sensitivity 
and thermal time constant. They had also (see Chap¬ 
ter 5) worked out a spring mounting to keep the bolom¬ 
eter strips under constant tension (see Chapter 3). 
This mounting had the effect of reducing micro¬ 
phonics. In order to embody the elements of design 
developed by MIT into a structure suitable for mass 
production, the Division, with the cooperation of the 
Vacuum Tube Development Committee, made a 
contract with the General Electric Company (OEM- 
sr-1348). Under the provisions of this contract the 




226 


TRANSITION AND ENGINEERING ACTIVITIES 


company revised the design of the Heat Research 
Laboratory, applying the well-known techniques of 
vacuum-tube manufacturing. These designs were 
used by the subcontractors just mentioned in manu¬ 
facturing bolometers for field use. The Division made 
contracts with the Fairchild Camera and Instrument 
Corporation for complete scanning units and with 
the General Instrument Company, in Newark, New 
Jersey, for scanner amplifiers. The coordination of all 
of these contracts with the activities of the Heat 
Research Laboratory at MIT (see Chapter 3) was 
carried on almost entirely by the Engineering and 
Transition Office. Personnel who were made avail¬ 
able by them to the Division for full-time work 
undertook activities ranging all the way from super¬ 
vising field tests to acting as purchasing and procure¬ 
ment agents for the Division contractors and their 
subcontractors. 

Through the contract of the NDRC Transition 
Office with the New England Power Service Com¬ 
pany (OEMsr-1260), personnel in several specialized 
fields of engineering were made available. These 
engineers undertook the construction of targets at 
proving grounds from designs developed by the Heat 
Research Laboratory. They made a special trip to an 
arsenal and there designed and built special tools to 
measure the variation in standard bomb dimensions 
from approved drawings, allowable tolerances being 
unavailable from the Ordnance Department. They 
developed a technique for applying supporting bands 
to the bomb body, using patch bolts with heads 
which sheared off at a known loading; thus it was 
impossible to prestress the holding bolt or supporting 
band so as to impair the factor of safety against 
stresses due to impact during landing of an aircraft 
with a bomb load. They sought out materials im¬ 
pervious to tropical fungi to replace felt gaskets used 
by the laboratories in early work, and organized and 
systematized the production layout. They checked 
many hundred production drawings to ensure that 
the pilot models would operate in accordance with 
the scientific principles developed by the MIT Heat 
Research Laboratory. 

Finally, one of the engineers from the NDRC 
Transition Office continually supported the Division 
by maintaining a constant liaison between the re¬ 
search laboratory at MIT, the production contrac¬ 
tors, and the proving ground so that difficulties as 
they arose in field tests were promptly referred to the 
laboratory, corrected, and the correction embodied 
in the production design without delay. 


116 ROC AND ITS CONTROLS 

The development program of Roc and its controls 
never reached the stage of production for combat. 
The situation with respect to this project, however, 
is sufficiently different from the normal run of a proj¬ 
ect as carried forward in the Division to its final 
stage to deserve special mention. 

The development agencies were each commercial 
production units: Douglas Aircraft Company for the 
missile and its servo links, and Bendix Aviation, 
Limited for the electronic control equipment. Engi¬ 
neering details, therefore, as well as the development 
of scientific principles, were assigned to them by the 
Division. The Division headquarters maintained a 
supervisory responsibility for both contractors and 
undertook to coordinate their activities. There can 
be no question that engineering for production is bet¬ 
ter accomplished by commercial organizations than 
by governmental or endowed laboratories, whether 
devoted solely to research or to research and educa¬ 
tion. A danger, however, exists in lodging both re¬ 
search and engineering phases of a program in a 
single contractor. Unless strong supervision is main¬ 
tained it is easily possible for the Government to lose 
control of the program. In times of peace this might, 
perhaps, not be serious, since the personnel of the 
contractor is almost certainly more competent in 
detail than any personnel the Government is likely 
to assemble. In time of war, however, this is not 
necessarily true. Competent industrial organizations 
during both World Wars have been heavily loaded 
with war work, and those projects not operating 
under strict supervision are likely to be assigned to 
the least competent staff of the contractor’s organi¬ 
zation. 

This is not to say that the Division experienced 
this evil in either of the contractors out of which the 
Roc program emerged. Rather it is a warning against 
the assumption by a Government agency that the 
need for strict supervision vanishes just because the 
contractor’s organization is prepared to undertake 
both basic research and reduction to practice. 

117 MISCELLANEOUS 

ENGINEERING ACTIVITIES 1 

In connection with all its guided-missile programs, 
the Division used consulting engineering services 
made available to it by MIT under Contract No. 
OEMsr-240. The activities covered by the engineers 




MISCELLANEOUS ENGINEERING ACTIVITIES 


assigned by MIT to this contract were extremely 
broad in character. 

11,71 Azon Program 

In connection with the Azon program, a study of 
storage batteries was made, resulting in the develop¬ 
ment by Willard Storage Battery Company of a 12-v, 
nominal 10-amp lead and acid storage battery. As a 
by-product of this program, the company was able to 
increase the capacity of the NT-6 battery used in 
Azon so that it was adequate to handle the increased 
load of the 2,000-lb Azon and of Razon. 

Increased storage battery capacity was not ob¬ 
tained immediately. Considerable work had to be 
done by the Willard Storage Battery Company; as a 
parallel enterprise the project engineers for the Divi¬ 
sion requested the Division contractor responsible 
for the missile to redesign the components so that the 
current drain was materially reduced. Through the 
combination of these activities the reliability of the 
missile in long missions, where low temperatures can 
be expected, was materially increased. 

The 2,000-lb Azon VB-2 was engineered for pro¬ 
duction almost wholly under the supervision of the 
engineers made available by MIT. They worked 
closely between Division headquarters and the Divi¬ 
sion’s contractor, the Union Switch and Signal Com¬ 
pany. When pilot units became available they super¬ 
vised drop tests at the Army Air Base at Tonopah, 
Nevada, reporting continually to the Division. Fi¬ 
nally, they cooperated with the Army Air Forces in 
the performance of evaluation tests by the Army Air 
Forces Board and assisted in analyzing the data re¬ 
sulting from this testing program. 

11,7,2 Razon Program 

The activity of these engineers in connection with 
the Razon program was identical with the service 
performed by them in connection with Azon. In addi¬ 
tion to supervising the production engineering on 


227 


these units by the Union Switch and Signal Company 
under Contract No. OEMsr-1415, they performed 
proof drops at Wendover Army Air Base with the 
cooperation of the Air Technical Service Command 
and supplied personnel to cooperate with the Army 
Air Forces Board for evaluation tests at Orlando, 
Florida. 

The Guided Missiles Subcommittee of the Joint 
Committee on New Weapon^ and Equipment of the 
United States Joint Chiefs of Staff had recommended 
to the Assistant Chief of Air Staff that the services of 
NDRC be utilized in connection with the evaluation 
program at Orlando. Personnel from the Gulf Lab¬ 
oratory instructed Air Forces personnel in the use of 
the weapon. Personnel from the Applied Mathema¬ 
tics Panel assisted in planning the tests and analyzed 
their results. The coordination of the activities of 
both these groups with the activities of the military 
was handled by the MIT engineers. 


11,7,3 Additional Engineering Activities 

In addition to carrying a considerable responsi¬ 
bility in the two major programs just mentioned, 
these engineers were continually occupied with a 
large number of miscellaneous assignments. A few of 
them were: expediting the production of fifty Crab 
attachments to be made available for experimental 
operational use in the Pacific Theater; the search for 
more reliable flares of higher intensity and studies of 
color discrimination with filters for multiple-release 
operation; the investigation of fuze drives which 
would not impair the aerodynamic performance of 
the missile; the elimination of radio noise produced 
by gyroscope motors; supervision of the procurement 
of bombing trainers for guided missiles (see Chapter 
10 ). 

As World War II closed, the representative of this 
group had been alerted for overseas service as a 
scientific consultant to the Thirteenth Air Force. The 
cessation of hostilities made this mission unnecessary. 










PART III 

CONTRIBUTIONS FROM INDIVIDUAL INVESTIGATORS 









Chapter 12 

SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


12.1 INTRODUCTION 

M odern warfare is primarily a warfare of missiles 
—bullets, shells, rockets, bombs, and grenades. 
The chief problems are to transport, project or pro¬ 
pel, and direct these missiles accurately to strike tar¬ 
gets of military importance. Missiles may be trans¬ 
ported to the vicinity of a target by men on foot, by 
mobile artillery, by tanks, by aircraft, or by ships. 
They may be dropped from aircraft, projected from 
guns on land, ship, or aircraft, accelerated for all or 
parts of their trajectories by rocket motors, or they 
may be self-propelled. Most of the missiles in current 
use are not under the control of the user after launch¬ 
ing. Some form of sighting device is necessary to es¬ 
tablish the initial direction of travel in a manner to 
cause the missile to strike the target. Gunsights, 
bombsights, and complicated gun directors and com¬ 
puters are mere aids to determine where to point the 
gun or release the bomb, and after release of the 
missile no further control is possible. 

For many years much thought has been given to 
the possibilities of guiding missiles after their release 
to correct for errors in sighting and for evasive action 
of moving targets. The development of radio com¬ 
munication stimulated inventive activity in this field 
and even during World War I there were experi¬ 
ments on radio-directed aircraft. There are innumer¬ 
able patent applications relating not only to radio 
control but also to other aspects of guided missiles. 
Particularly attractive has been the idea of missiles 
which will seek out a target automatically or “home” 
on the target. 

During the present war much serious scientific ef¬ 
fort has been devoted to a full exploration of the 
practicability of guided missiles as useful weapons. 
The German experiments with Hs 293 and FX and 
later with V-l and V-2 stimulated the interest of our 
military leaders and enlisted their support in the fullest 
exploitation of the possibilities of this type of weapon. 

1211 Classification of Missiles 

From the standpoint of the source of control, there 
are five possible classes of missile. The first, illus¬ 
trated by the German V-l and V-2, uses an autopilot 


to hold the missile on a prescribed course. The missile 
receives no information relating to the target, and 
must be launched in the direction required to strike 
the target. This type is really a kind of long-range 
artillery, although the missile itself carries a part of 
the computer. 

The second type, illustrated by the Japanese sui¬ 
cide bombers, carries a human pilot in the missile 
who is expended with the missile. This solution is 
known, involves no new technical problems, but is 
not attractive to a nation placing a high value on 
human life. 

The third type, illustrated by the German Hs 293 
and FX and the Division 5 NDRC Azon and Razon 
series, uses a remote human pilot who may (1) see 
the missile and target visually, aided by flares on the 
missile and by sighting devices, (2) be guided by tele¬ 
vision or radar information furnished him by ap¬ 
paratus contained in the missile, or (3) be guided by 
radar location of the missile in relation to map or 
radar location of the target. 

The fourth type, of which there is no existing il¬ 
lustration, uses some type of beam directed toward 
the target which the missile automatically follows. 

The fifth type is the target-seeking or homing type, 
illustrated by the Division 5 NDRC Pelican, Bat, 
and Felix. Such a missile must utilize some physical 
property of the military target which causes it to 
stand out from the background. The most commonly 
suggested property is the emission or reflection of 
electromagnetic radiation. Separate techniques are 
available for transforming three major divisions of 
the electromagnetic spectrum into directional infor¬ 
mation. Of the three—visible light, infrared, and 
radio—radio frequencies, for technical reasons, hold 
the most promise for useful weapons. Radio tech¬ 
niques as developed in radar are directly applicable. 
Visible light and infrared are useful for certain spe¬ 
cific types of targets. Investigations in other suggested 
properties, such as the emission or reflection of sound, 
have not given promising results for aero-missiles. 

121,2 Necessity for Technical Coordination 

This chapter is concerned with various aspects of 
the design of this last type, the homing aero-missile. 


231 


232 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


The impression is prevalent that scientific advances 
in many fields have progressed to the point where the 
development of such a missile is purely a matter of 
engineering design on the part of the several special¬ 
ist groups, with the usual coordination as to dimen¬ 
sional requirements, weights, and time of completion. 
Experience has taught otherwise. Optimistic time 
schedules based on such an assumption cannot be 
met. The development of successful homing aero- 
missiles requires the solution of certain research prob¬ 
lems associated with the complete device, involving 
complex relationships between the performance char¬ 
acteristics of the component parts. There is required 
a type of overall technical coordination beyond that 
required in the design of aircraft as ordinarily prac¬ 
ticed. 

The necessary technical coordination is made diffi¬ 
cult by the wide variety of specialists required, each 
with different scientific and technical backgrounds 
and accustomed to different vocabularies and habits 
of thought. For example, in the case of a propelled 
radar-homing missile there will be represented ex¬ 
perts in aerodynamics, aircraft structures, propul¬ 
sion, servomechanisms, electronics, radar, computers, 
explosives, and fuzes. Other types of missiles will re¬ 
quire experts in radio, optics, infrared radiation, heat, 
etc. No one person can be expert in all these diverse 
fields, but the success of the project requires a project 
engineer who has sufficient knowledge of these fields 
to be able, with the help of advice from the special¬ 
ists, to assume technical leadership in the solution of 
research problems associated with the system as a 
whole. 

In the design of any homing missile, there soon 
emerges a number of problems which cut across the 
boundaries of the specialist groups. The particular 
design which seems best to one group of specialists 
creates difficult problems for other groups, and the 
requirements put forth by the several groups as op¬ 
timum are often contradictory. For example, certain 
errors are introduced unless the intelligence device 
“looks” along the direction of motion, i.e., is accu¬ 
rately bore-sighted. The conventional airplane flies 
at an angle of attack which varies with the position of 
the elevator. If the aerodynamics specialist adopts a 
conventional aircraft design with elevator control, 
the intelligence device must be coupled to the eleva¬ 
tor control in such a manner as to compensate for 
variations of angle of attack for all conditions of 
flight. However, the designer of the intelligence de¬ 
vice might properly suggest that the aerodynamicist 


design a vehicle which does not change its attitude 
with application of the controls. It becomes a 
matter of research to determine which solution gives 
greater accuracy and hence how responsibility is to 
be allocated between the two specialists. 

121,3 Purpose and Background 

These broad aspects of the design of homing aero- 
missiles are treated in this chapter. An attempt has 
been made to make the discussion general in charac¬ 
ter and applicable to all such missiles, whether pro¬ 
pelled or not. It should be stated, however, that the 
discussion arises from experience with the radar¬ 
homing missiles of the Pelican and Bat series, and 
this account undoubtedly reflects the solutions there 
adopted, as well as the problems peculiar to radar 
homing. 

12 2 TARGET DISCRIMINATION 
AND TRACKING 

12,2,1 Limitations of Mechanisms 

In visual shooting or bombing, the target is identi¬ 
fied by the pattern of optical radiation as perceived 
by the human eye and interpreted by the human 
brain. In radar fire control or bombing, the target is 
identified by the pattern of short-wave electromag¬ 
netic radiation exhibited on the screen of a cathode- 
ray tube as perceived by the human eye and inter¬ 
preted by the human brain. During the flight of a 
homing aero-missile these radiation patterns must be 
made to operate control mechanisms, and the ele¬ 
ment of interpretation by a human brain is absent. 
This introduces many problems and severe limita¬ 
tions. A mechanism can perceive only a limited num¬ 
ber of physical characteristics of the target pattern 
and can exercise no judgment in interpreting the 
received information except of the simplest kind— 
for example, smoothing over a certain time interval. 
At present no mechanism is known which will select a 
given target from an optical image of a complex vis¬ 
ual pattern. In their present state of development 
homing missiles can utilize only simple target situa¬ 
tions, the most favorable target situation being that 
of a small number of ships on a large body of water 
or, more generally, a few isolated discontinuities in a 
radiation pattern which is otherwise nearly uniform. 
From this point of view, aircraft also present favor¬ 
able target situations. 



TARGET DISCRIMINATION AND TRACKING 


233 


12.2.2 p r i nc ipl es Q f Target Discrimination 

Except in the case, which seldom occurs, of a single 
isolated discontinuity in the radiation pattern, the 
first main problem in designing any homing missile is 
to determine the method or methods to be used for 
target discrimination when several targets are pres¬ 
ent. In the simpler devices the missile homes on the 
biggest discontinuity within its field of view. In many 
homing missiles, means are provided to permit the 
initial selection of the target by a human operator 
who often has available information supplemental to 
that given to the control circuits of the missile. 
Means may also be provided within the missile to 
keep track of the initially selected target and to re¬ 
spond only to signals from that target. Even in the 
simpler missiles, the operator attempts to release the 
missile either with a single large discontinuity within 
the field or on such a path that a single target comes 
within the field, and he depends on the precision of 
the controls of the missile to keep the missile track¬ 
ing this target. 

In order to obtain information about the location 
of the target, a homing missile must be directionally 
sensitive. Usually the received intensity of radiation 
is greatest when the axis of the radiation-receiver is 
pointed directly at the target. Figure 1 shows a por¬ 
tion of the response curve of the receiving antenna of 
the radar receiver used in the Pelican project. The 
relative power is plotted in terms of decibels, a loga¬ 
rithmic unit, but the power ratios are also indicated. 
The width at l U power is 23 degrees, and the width 
at Vio power is 43 degrees. In other words a target at a 
bearing 2iy 2 degrees from the antenna axis gives ^10 
as much energy to the receiver as a target on the axis 
giving the same intensity of radiation. 

The radiation pattern shown cannot be used 
directly because there is no discrimination between 
right and left or up and down. The usual practice is 
to scan the field of view, to commutate or phase the 
received signal intensity with the scanning, and com¬ 
pare right with left and up with down or perform 
comparisons in some other coordinate system. The 
on-course indication becomes then an equality of two 
signals, and a directional sense is provided. In Peli¬ 
can, conical scanning is used, the antenna axis des¬ 
cribing a cone with a half-vertex angle of 11 degrees. 
A commutator provides the phasing. If there were a 
variety of targets of equal intensity, every one within 
a cone of half-vertex angle of approximately 22 de¬ 
grees would give signals of one-half maximum power 


or more. This represents one method of stating a 
figure for the field of view of the receiver, all targets 
being assumed to return signals of equal intensity. 

In Pelican, the target is illuminated by a radar 
beam, and the directional characteristics of the trans¬ 
mitter antenna provide additional directional dis¬ 
crimination, which is not of interest in this discus¬ 
sion. In Bat, the missile carries the transmitter and 
the same antenna is used for transmission and recep¬ 
tion. The field of view is accordingly smaller than for 
Pelican. 

In other types of intelligence devices, much smaller 
fields of view are used, the width at V 2 power being 
10 degrees or less. This is advantageous from the 
point of view of directional discrimination of targets. 
Limitation of the field of view is the first general 
method of securing target discrimination. However, 



Figure 1 . Directional sensitivity of antenna of Pelican 
radar receiver. 


a narrow field of view introduces tracking problems, 
as will be discussed later. 

The second method of securing target discrimina¬ 
tion is by means of signal strength. This cannot be 
entirely separated from the directional properties of 
the intelligence system and permits little choice other 
than to home on the strongest signal. It is usually 
necessary to include some type of automatic gain 
control to obtain directional information at signal 
levels which may vary more than a billionfold as the 
missile approaches the target. The strongest signal 
within the field of view will govern the sensitivity of 
the receiver through the action of the automatic gain 
controller. 

Target discrimination may also be secured through 
the selective action of the intelligence device in re¬ 
sponding to radiation within narrow wavelength 
limits. This is best utilized when an intense beam of 
the desired radiation can be concentrated on the tar- 





















234 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


get and the missile made sensitive only to wave¬ 
lengths within the narrow limits of the transmitted 
radiation. 

In systems in which a pulsed illuminating beam is 
used, as exemplified particularly in the Pelican and 
Bat radar-homing systems, another method of target 
discrimination may be used. This is range selection 
and synchronization. By making the intelligence sys¬ 
tem sensitive only for a short period at a predeter¬ 
mined time following the emission of an illuminating 
pulse, the control information can be restricted to 
that received from targets lying within certain range 
limits, say within a zone of ± 250 ft of the actual target 
range. A given range corresponds to a definite time of 
transit of the pulse from transmitter to target to re¬ 
ceiver. Naturally, such a range selector requires an 
automatic method of tracking the target in range as 
the range decreases. The operation of the range selec¬ 
tor also requires the synchronization of the receiver 
and the transmitter. The synchronization is effective 
in discriminating against reflected radiation originat¬ 
ing from other transmitters operating on the same 
radio frequency but with different pulse rates. 

The use of range selection is found to be essential 
in radar-homing missiles launched from aircraft be¬ 
cause of the so-called altitude signal , i.e., energy re¬ 
turned from the earth directly below the aircraft. If a 
reasonable cone of vision is to be maintained, the 
directional selectivity of the antenna is insufficient to 
discriminate against the large reflecting area lying 
beneath the aircraft. A range selector and range¬ 
tracking device makes possible the elimination of this 
signal in the case of glider missiles, since the altitude 
is always less than the range to the target. Presently 
available radar-homing devices cannot be used under 
conditions where the target may appear at the same 
range as the altitude signal—for example, in air-to- 
air missiles at ranges greater than the altitude. In the 
case of ground-to-air missiles using only a receiver, 
the geometry is more favorable than for the glider, 
and the range of the target will not coincide with that 
of the altitude signal. For a send-receive missile, the 
missile will at some time be at the same altitude as 
the range to the target and hence may thereafter 
home on the altitude signal. 

It is possible to devise more complex mechanisms 
to perform more difficult feats of discrimination, for 
example, to permit the launching of a radar-homing 
missile at long range without advance selection of a 
target and to have the range-selection device search 
for, choose, and lock on a target when the missile has 


proceeded a definite distance. The limit of perform¬ 
ance is set only by the permissible complexity of the 
mechanism. 

The suggestion has often been made that lower ani¬ 
mals be used as intelligence devices, since their 
brains, like human ones, can perform difficult tasks of 
discrimination. This possibility is perhaps the only 
one of adequate complexity to deal with the pattern 
discrimination required to select, for example, a 
particular building within the complex optical radia¬ 
tion pattern presented by a city. Proponents of this 
method point out that almost mechanical reliability 
may be realized in animals by establishing in them a 
conditioned reflex associated with the object selected 
for attention. 

12,2,3 Relation between Field of View 
and Permissible Motion of Vehicle 

After a target has been selected by the operator 
before release of the missile or by the mechanism of 
the missile itself, the target must be tracked, i.e., 
held within the field of view of the missile during the 
remainder of its flight. The simplest method of track¬ 
ing is to have the intelligence control the motion of 
the missile so that the target remains within the field. 
If this method of tracking is selected, restrictions are 
immediately placed on the permissible motions of the 
vehicle, which must be considered by the aerody¬ 
namics specialist. These restrictions depend not only 
on the aerodynamics of the vehicle and the field of 
view of the intelligence device but also on target con¬ 
trast, characteristics of tracking circuits, and be¬ 
havior of the servomechanism in the absence of hom¬ 
ing signals. 

When a missile is to be released blind, the aero¬ 
dynamics specialist can compute trajectories for vari¬ 
ous release conditions and so provide estimates of the 
time at which a target in a specific location relative to 
the point of release will lie within a specified field of 
view. The relation between field of view, servomech¬ 
anism, and aerodynamical characteristics must be 
such that tracking will be preserved. The controls 
must be sufficiently effective to check any overshoot, 
or the servomechanism must have a memory to bring 
the vehicle and field of view of the intelligence device 
back on the target. For some types of vehicles, the 
aerodynamic design can be made such that the tra¬ 
jectory with no homing signal will include the de¬ 
sired target within the field of view. The larger the 
field of view, the easier the task. 






TARGET DISCRIMINATION AND TRACKING 


235 


Before release of the vehicle, the target must be 
brought into its field of view and tracking in range 
established. During launching and thereafter, the 
motion of the vehicle must be such as to maintain the 
target within the field of view of the intelligence de¬ 
vice. Memory circuits within the intelligence head 
may allow the signal to fade out for short time inter¬ 
vals and resume tracking in range when the signal 
level is restored. If during the period of no signal 
return, the vehicle motion removes the target from 
the field of view, the intelligence device is unable to 
secure further information on target direction. The 
permissible motion will depend upon the intensity of 
the signal returned by the target. If the target signal 
is only a small amount of the background signal, the 
effective cone of vision is reduced (in present radar¬ 
homing equipment) to about 70 per cent of its maxi¬ 
mum width. Thus a smaller change in attitude will 
be required to lose directional tracking than if the 
target signal is much larger than the background. 
Since the average signal level and its consistency in 
amplitude depends on target size, target orientation, 
and meteorological conditions, it is difficult to give 
definite design rules. However, the smaller the field of 
view, the smaller the change in attitude required to 
reduce a low signal to the background level, and from 
these considerations a large field of view is desirable. 

The behavior of the servomechanism in the ab¬ 
sence of the signal, or, more exactly, just following 
fading of the signal, has a definite bearing on the rela¬ 
tion between field of view and permissible motions of 
the vehicle. If the servomechanism maintains the 
vehicle on the course it was flying, the field of view 
can be small without risk of losing the target outside 
this field should the signal fade for a few seconds. If, 
however, the servomechanism maintains the rates of 
turn and pitch which exist at the time of signal fad¬ 
ing, the target would probably pass outside a small 
field of view before the signal returned. Either of 
these types of performance of a servomechanism is 
somewhat idealized and not accurately obtainable 
in any actual mechanism. The maintenance of the 
same course is of advantage when the vehicle is 
nearly on the desired course and loss of signal is due 
to fading. 

When the vehicle is initially coming on course, it 
may overshoot by a sufficient amount to lose track. 
If this occurs, a servomechanism which maintains the 
course of the vehicle at the time the signal is lost will 
thereafter give no opportunity for again picking up 
the target signal. The amount of overshoot permis¬ 


sible will be larger, the larger the field of view and the 
greater the target contrast. 

Some of the restrictions on the motion of the ve¬ 
hicle which are imposed by a narrow field of view can 
be removed by the use of an intelligence device fitted 
with automatic directional tracking. In this system 
the output of the intelligence device is used to drive 
servomotors to center the field of view on the target 
independent of the motion of the vehicle. The control 
of the vehicle itself is then derived from the relative 
position or rate of motion of the intelligence device 
with respect to the vehicle, or both. The minimum 
permissible field of view is then limited only by the 
precision and speed of response of the servomechan¬ 
ism. The extra degrees of freedom may give rise to 
more difficulty with the stability of the two servo¬ 
mechanisms—one driving the intelligence, the other 
the vehicle. There has, as yet, been no field exper¬ 
ience with a missile control of this type. 

12 2 4 Background Signal 

All electronic intelligence devices have a certain 
internal noise level which cannot be less than that 
produced by thermal agitation of electrons in the in¬ 
put circuit. The magnitude of this internal noise sets 
a lower limit to the signal, which can be detected. 
However, in actual practice there is a much higher 
background signal, representing the signal return 
from areas other than that of the target which also 
lie within the field of view. Thus in a radar-homing 
device, the background signal is the reflected radar 
energy from the land, rough sea, or other obstruc¬ 
tions that happen to be at the same range as the tar¬ 
get. It may be very small or zero when the target is 
an airplane and the background is cloudless sky. In 
an optical homing device, the background signal is 
the reflected or emitted optical energy from land, sea, 
or sky. 

The important attribute of this unwanted back¬ 
ground energy is its variability, not only from place 
to place but also at the same place, especially with 
weather. Where radar is used against ship targets, 
the “sea return” depends very greatly on the height 
and shape of the waves and on the orientation of the 
wave troughs with respect to the receiver. The per¬ 
formance of a radar-homing missile against a speci¬ 
fied target is affected very much by the condition of 
the sea surface, the permissible range at release being 
reduced as the sea becomes heavier because the sig¬ 
nal returned by the ship is smaller and the amplitude 



236 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


of signal fluctuation is greater. Small targets, regard¬ 
less of their range, may be lost in the sea return. 
Similarly, missiles using other parts of the electro¬ 
magnetic spectrum encounter background signals 
which usually depend greatly on meteorological 
conditions. 

The ratio of target to background signal may show 
large short-time variations during the flight of a sin¬ 
gle missile, and if the target contrast is not sufficient, 
tracking at long range may be difficult. 


12 2 5 Fluctuation of Signal Intensity 

As a homing missile approaches its target, the sig¬ 
nal intensity and, usually, the background signals 
increase very greatly, making necessary an auto¬ 
matic gain control in the electronic equipment. The 
time constant of the gain control must be short 
enough to take care of the rapid change at the end of 
the flight but not so short as to obscure the variations 
of scanning frequency which give the directional in¬ 
formation. The reliability of the information obtain¬ 
able is dependent on the target contrast, which is a 
function of the strength of both the background sig¬ 
nal and the target signal itself. In addition to the 
slower variation of signal strength as the missile ap¬ 
proaches the target, there may also be more rapid 
variations associated with the changing geometrical 
relationship between missile and target produced by 
the motion of the missile and the linear and angular 
motions of the target. Such fluctuations are always 
found in radar reflections. 

The presence of fluctuating signal intensity and 
fluctuating background signal means that the missile 
and servomechanism cannot be designed on the as¬ 
sumption that information as to the angular bearing 
of the target is continuously available. 

The effects of fluctuating signal intensity which 
must be guarded against are possible resonance ef¬ 
fects in the servomechanism, synchronization with 
the frequency of scanning, and loss of tracking. It is 
not practical to lay down methods of design. How¬ 
ever, in the radar case the MIT field experiment 
group working on Pelican and Bat have found it ad¬ 
vantageous to make photographic records of varia¬ 
tions of signal strength of actual targets, to construct 
a special signal generator which emits variable sig¬ 
nals controlled by a cam cut to the observed varia¬ 
tions, and to test the effects of such a signal input on 
the intelligence device output. 


123 MANEUVERABILITY OF VEHICLE 
AND OTHER AERODYNAMIC PROBLEMS 

12,31 The Six Degrees of Freedom 

The trajectory of a missile is determined by the 
force of gravity and the reactions between the missile 
and the air through which it flies. The force of grav¬ 
ity acts vertically downward through the center of 
gravity of the missile. It is convenient to represent 
the resultant of the aerodynamic reactions by three 
mutually perpendicular force components acting- 
through the center of gravity and three moments act¬ 
ing about the three axes along which the force com¬ 
ponents are taken. These forces and moments on a 
given missile are functions of the shape of the missile, 
the orientation relative to the direction of motion of 
its center of gravity, the speed, the axis and amount 
of angular rotation, and the density and other physi¬ 
cal properties of the air. 

The missile has six degrees of freedom, three linear 
and three angular. The user of a missile is interested 
essentially in the three linear degrees of freedom, i.e., 
in the linear motion of the center of gravity of the 
missile. The angular motion of the missile is of inter¬ 
est only as it modifies the three force components and 
thus the trajectory. A spinning or angular hunting- 
motion is of no interest if the missile strikes the tar¬ 
get, a result dependent only on the path of the cen¬ 
ter of gravity. It is generally true for the ordinary 
bomb, propelled aircraft, or glider that an absence of 
angular motion gives a more predictable and con¬ 
stant trajectory. But in some other missiles (for ex¬ 
ample, shells), a spinning motion is deliberately pro¬ 
vided to give a more stable and predictable trajec¬ 
tory. 

The dynamics of a missile, even if restricted to air- 
plane-like or bomb-like objects, is a very complex 
subject and hardly appropriate for this chapter. 
(Readers are referred to Aerodynamic Theory 1 for a 
discussion of airplane dynamics.) Only elementary 
and general aspects will be discussed here. 

1 he path of the missile can be controlled in a num¬ 
ber of ways, but the most usual method is through 
changes in the aerodynamic reactions by means of 
changes in the shape of the missile. Other methods 
may be advantageous in special cases. Thus, the use 
of rocket motors makes it possible to apply reaction 
forces on the missile to change its path. This method 
is effective in the stratosphere, where the air density 
is very small, and in free space. Control requires the 
ejection of a part of the missile, a process differing 




MANEUVERABILITY OF VEHICLE AND OTHER AERODYNAMIC PROBLEMS 


237 


only in degree from the burning of fuel to produce the 
power that operates other types of control devices. 
Where other types of control are possible, it is usu¬ 
ally more economical to use them. 

A missile may be controlled by varying its mass 
distribution to modify the position of the center of 
gravity, thus changing the resultant moments of the 
air reaction—hence the orientation of the missile, 
which in turn modifies the force components. This 
was done in early airplanes by motion of the pilot but 
the method has been little used since that time. 

It is possible to use power-driven devices, such as 
propellers or turbojets, or to use thermal jets to pro¬ 
duce forces that modify the path, or to use such de¬ 
vices to apply moments to the missile that change its 
orientation and thus produce forces to modify the 
path. It is more common to use these devices as pro¬ 
pulsion elements, and some control of the path, es¬ 
pecially changes in the vertical plane, is accomplished 
by varying the propulsion force. 

The most common method of control is through 
changes in shape of the missile which usually alter 
the moments of the air reaction and change the orien¬ 
tation of the missile in addition to modifying the 
force components directly. The use of a power-driven 
propeller may be regarded as a special case of a per¬ 
iodic change in shape. All changes of shape for pur¬ 
poses of control involve the application of power, 
which may be derived from any of the usual types of 
power sources. It is obviously desirable that the 
power required for control be small. This has led to 
the conventional type of control used on aircraft, in 
which the primary effect of the controls is to apply 
moments to the missile, which, in a time determined 
by the angular moments of inertia and the magnitude 
of the applied moments, changes the orientation of 
the missile to the direction of motion of its center of 
gravity. As a result of the change in orientation, the 
forces are changed and the path of the center of grav¬ 
ity is modified. It will be seen later that there are 
advantages, when designing homing aero-missiles, in 
selecting a change in shape which produces little or 
no change in moment but does produce directly 
changes in the force components. Such a method, 
however, requires greater power for operating the 
controls than the conventional method. 

12 3 2 Equilibrium and Trim 

An unpropelled missile can be in complete equilib¬ 
rium only if the moments of the air reaction about 


three mutually perpendicular axes through its center 
of gravity total zero and if the resultant air force is 
equal to the weight of the missile and acts vertically 
upward. Such a state does not exclude the possibility 
of a spiral or spinning motion. In fact the tailspin of 
an aircraft is a steady motion in which the above con¬ 
ditions are fulfilled. Such motions, however, will not 
be considered further. Although there is no logical 
necessity of excluding spinning missiles, the guiding 
of such missiles would seem to introduce many tech¬ 
nical complications. 

The equilibrium of a propelled missile differs only 
in that the resultant aerodynamic force must equal 
the resultant of the weight and the propelling force. 

Practically all missiles now used or under consider¬ 
ation have one or two planes of symmetry and a lon¬ 
gitudinal axis which lies within 10 degrees of the in¬ 
tended direction of motion. The exact location of the 
longitudinal axis is usually chosen to suit the con¬ 
venience of the specific problems but always in a 
plane of symmetry. The other mutually perpendicu¬ 
lar reference axes are called the lateral and normal 
axes, and if the missile has only one plane of sym¬ 
metry, this plane contains the longitudinal and nor¬ 
mal axes. In the practical construction of aircraft or 
missiles it is found impossible to make the device suf¬ 
ficiently accurately to ensure that, when flown or 
released, the aerodynamic moments about the refer¬ 
ence axes will be zero. It is always necessary to apply 
control moments of suitable magnitude about all 
three axes, or if it is desired to have the controls in a 
given neutrai position with no force applied to the 
control levers, to provide adjustable trim tabs. These 
adjustments are easily made when a human pilot is 
on board, but other provisions must be made when 
unmanned missiles are to be used. 

For unbalanced moments about the lateral and 
normal axes, a stable missile compensates by angular 
rotation to new positions of equilibrium, since dis¬ 
placements about these axes produce restoring mo¬ 
ments. The missile would then fly at a somewhat 
different angle of attack than planned and at an 
angle of yaw which would give rise to a lateral force 
producing a lateral drift of the missile. An unbal¬ 
anced moment about the longitudinal axis, which lies 
approximately in the direction of motion, cannot be 
compensated in this way, because rolling about the 
direction of motion produces no static restoring mo¬ 
ment. The only methods known of compensating for 
this unavoidable and undesired moment arising from 
lack of symmetry in the actual unmanned missiles 




238 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


involve automatic trimming by control-surface dis¬ 
placements under the control of one or more gyro¬ 
scopes. The most desirable method is to compensate 
for the undesired moment directly by displacement of 
the ailerons. In some cases, for example, the German 
robot bomb V-l, the autopilot moves the rudder, 
thus forcing the bomb to travel at an angle of yaw 
sufficient to produce a rolling moment due to yaw 
equal to the unbalanced rolling moment. This meth¬ 
od of correction gives rise to a lateral drift, which is 
one source of error contributing to the dispersion. 
One of the results of the early work on the Pelican 
and Bat developments was the demonstration that 
provision must be made in the autopilot for compen¬ 
sating for undesired aerodynamic moments arising 
from lack of symmetry, i.e., “trimming” the missile, 
and that a gyro or equivalent reference is essential. It 
is fairly well known that pendulums or aerodynamic 
surfaces whose position is controlled by aerodynamic 
reactions are ineffective for this purpose. In the Azon 
and Razon developments it has also been found de¬ 
sirable to introduce automatic trim devices to elim¬ 
inate rolling motions, although not there required for 
stability reasons, since these missiles have two planes 
of symmetry. The elimination of the rolling motion 
in those missiles simplifies the control problem, as 
will be discussed in the section on coupling between 
controls. 

The state ol complete equilibrium is rarely at¬ 
tained in the practical use of missiles. The linear mo¬ 
tion of the center of gravity is an accelerated one, the 
mass times the acceleration being equal to and in the 
direction of the resultant force. The time required to 
reach equilibrium is often much greater than the time 
of flight of the missile. For example, a falling bomb is 
in complete equilibrium only when it reaches its ter¬ 
minal velocity, a process requiring fall from a great 
height and many tens of seconds. This long time con¬ 
stant arises from the limited rate at which energy is 
supplied from the gravitational field. In the case of an 
aircraft the slow phugoid oscillation arising from in¬ 
terchange of kinetic energy and potential energy has 
a period of approximately 0.22V seconds when V is 
the speed in ft per sec, i.e., the period is 88 seconds 
for a missile traveling at 400 ft per sec. The design 
of servomechanisms and intelligence devices cannot 
be based on the assumption of equilibrium condi¬ 
tions. 

Fortunately the time constants of the angular mo¬ 
tions are much shorter, usually of the order of a frac¬ 
tion of a second or at the most a few seconds, increas¬ 


ing somewhat with the size of missile but decreasing 
with its speed. 

12 3 3 Magnitude of Lateral Forces, 

Angular Rates, and Radii of Curvature 

Consider a symmetrical missile falling vertically 
with its longitudinal axis also vertical. The forces 
acting are the force of gravity and the air resistance. 
Because of symmetry there are no lateral forces. The 
downward acceleration will be equal to the difference 
between the acceleration of gravity and the ratio of 
the air resistance to the mass of the missile. The ac¬ 
celeration will ultimately approach zero as the mis¬ 
sile approaches its terminal velocity, at which the air 
resistance equals the weight. At any point along its 
length the trajectory may be modified only by intro¬ 
ducing a lateral force. This force imparts a lateral 
acceleration which causes the missile to travel in a 
path which is approximately circular for some time. 
To move the missile in a path of radius R requires a 
centripetal acceleration of V 2 /R where V is the ve¬ 
locity of the missile. The rate of change of direction 
of the trajectory dQ/dt equals V/R. 

The usual method of securing a lateral force is to 
change the orientation of the missile so that its axis 
makes an angle to the trajectory. The missile does 
not then travel in the direction of its axis. The change 
in its direction of motion is dependent on the magni¬ 
tude of the lateral force produced by the change in 
orientation in relation to the mass of the missile. If 
the force is small or the mass is large, the trajectory 
will be modified very slowly, even though the axis of 
the missile is at a large angle to the direction of mo¬ 
tion of the center of gravity. The missile behaves in 
the same manner as an automobile traveling on ice 
when the steering wheel is suddenly turned. 

Experience from tests on bombs and airfoils shows 
that the lateral force produced at a given angle of the 
missile to its trajectory is approximately propor¬ 
tional to the square of the speed and to the projected 
lateral area. A reasonable value of the lateral force 
is about 10 psf on the fins or wings at a speed of 100 
ft per sec and about 3^ psf on a body of revolution 
at a speed of 100 ft per sec for angles of yaw of 15 
degrees, although with larger angles of yaw still 
higher values can be obtained. Extremely large an¬ 
gles of yaw give large drag forces which slow up the 
missile and thus reduce the lateral force. The lateral 
acceleration to be expected is therefore about 
(10A/W)(V/ 100) 2 g, where A is the area of the fins or 



MANEUVERABILITY OF VEHICLE AND OTHER AERODYNAMIC PROBLEMS 


239 


wings in sq ft, V the speed in ft per sec, W the weight 
of the missile in lb, and g is the acceleration of gravity 
in ft per sec per sec. In any actual design, the lateral 
acceleration should be determined from wind tunnel 
measurements on a model of the missile. A radius of 
curvature R requires an acceleration V 2 /R. Hence 


or 


0.0322 AV 2 


R 


W 


R = 31 


W 


The rate of change of direction dd/dt = AV/SIW. 

For the standard 2,000-lb bomb, A for the stand¬ 
ard fins is about 4 sq ft. With a suitable rudder, it 
may be expected that lateral forces of the above mag¬ 
nitude may be reached, in which case R = 15,500 ft 
and dO/dt = F/15,500 radians per sec = 0.0037 V 
degree per sec. In the next 500 ft of fall after the 
rudder is applied, the bomb would move laterally 
about 7ft. 

It is to be noted that for this case of a vertical tra¬ 
jectory, the radius of curvature obtainable with a 
given fin area does not depend on the speed, since 
both the required and the available acceleration vary 
as the square of the speed. However the forces which 
the missile structure must withstand increase in pro¬ 
portion to V 2 /R. 

When the trajectory makes an angle to the verti¬ 
cal, the force of gravity has a component at right an¬ 
gles to the trajecto^. It is customary to measure the 
angle from the horizontal and to denote its value by 
0. The gravity component is then g cos 0. The path 
then curves downward unless a sufficiently large lat¬ 
eral aerodynamic force overcomes the gravitational 
component. Calling the lateral aerodynamic force L, 
we have the following equation for the radius of cur¬ 
vature of the path. 


mV 2 dd _ nr 

~R~ = m '' (U = cos 0 ~ L 


Consider first the case in which L is zero, i.e., a 
conventional bomb. We find R = V 2 /{g cos 0) and 
dd/dt = (g cos d)/V. The maximum value of dd/dt 
occurs when the axis of the bomb is horizontal and 
equals g/V radians per sec. At a speed of 320 ft per 
sec, dd/dt = 0.1 radian per sec = 5.7 degrees per 
sec, and R = 3,200 ft. As the speed increases, dd/dt 
decreases and R increases. 

Next assume that the trajectory of the missile is 
to be approximately a straight line to the target, as 
is usually desired for a homing missile. In this case 


the average value of L must equal mg cos 0. The 
available control then depends on the changes in L 
which can be effected by the controls, and the control 
is the same as for a vertical trajectory. However, the 
requirement L = mg cos 0 is the requirement for rec¬ 
tilinear flight, and the speed required is dependent 
again on the area of the fins or wings. The minimum 
speed for rectilinear flight, assuming the use of the 
maximum control, is given by 



V_ = \/w cos_0 

100 V A 10 

For the 2,000-lb bomb on a 45-degree trajectory, V 
is about 600 ft per sec. 

At lower speeds than the minimum speed for recti¬ 
linear flight, the rate of change of direction is not as 
great, and the radius of tu^n is larger. Values for any 
particular case can be estimated. The control obtain¬ 
able depends greatly on the ratio of the speed to the 
rectilinear flying speed, and the rectilinear flying 
speed is determined mainly by the area of fins or 
wings. 


12 3 4 Nonlifting Missiles 

From the aerodynamic point of view the properties 
of missiles in which the average value of the lateral 
force is zero are so different from those for which the 
average value is not zero that the two groups deserve 
separate treatment. Although an aircraft or glider 
could be trimmed to give zero lift, its general flight 
behavior would not be satisfactory, and the nonlift¬ 
ing missile usually takes the form of an elongated 
object with two or more planes of symmetry and in 
some cases a body of revolution. The simplest ex¬ 
ample is an ordinary bomb or rocket stabilized by 
tail fins. When the axis is inclined, there is a restoring 
moment because the line of action of the resultant 
force passes behind the center of gravity. To keep the 
axis at an angle to provide a lateral force, this mo¬ 
ment must be balanced by a smaller force in the op¬ 
posite direction applied by a rudder at the tail or a 
spoiler at the nose. The action of the rudder or spoiler 
must be such that the missile still has sufficient static 
and dynamic stability about the new position of 
equilibrium. 

Missiles of this type are best adapted to steep tra¬ 
jectories for the following reasons. Launching speeds 
are usually limited to a few hundred feet per second. 








240 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


Control against the deflecting action of gravity is 
effective only near the rectilinear flying speed. The 
minimum rectilinear flying speed is of the order of 
100v / (TT/^4)(cos 0/10). As the path approaches the 
vertical, cos 0 decreases, the minimum rectilinear fly¬ 
ing speed decreases, and the control improves. 

A decrease in W/A also decreases the rectilinear 
flying speed and hence improves the control at low 
speeds. This principle has been used in the Roc proj¬ 
ect, in which additional surface has been provided to 
give larger lateral forces for a given angle to the tra¬ 
jectory. 

The control of a nonlifting missile in the horizontal 
plane is of course independent of gravity, and the 
general considerations discussed for a vertical tra¬ 
jectory apply. 

When control moments are applied to missiles of 
the nonlifting type about two axes at right angles to 
each other, the resultant air reaction cannot be rep¬ 
resented solely by a resultant force. There is also a 
resultant couple producing rotation about the longi¬ 
tudinal axis. In other words the missile has some of 
the characteristics of a screw propeller or windmill. 
It has usually been found desirable to provide ailer¬ 
ons controlled by gyro elements to keep the missile 
from rolling. Such a control is desirable to permit 
easy separation of right-left from up-down controls of 
the missile. 

It is inferred from the successful use of the German 
V-2 rocket, which is a nonlifting missile, that the 
problems of controlling missiles of this type at super¬ 
sonic speeds will not be too difficult. 


Lifting Missiles 

The nonlifting missile follows the normal type of 
approximately parabolic trajectory except when the 
controls are operated. The fins or wings ordinarily 
exert no lateral force. If, however, the vertical con¬ 
trol is held deflected so that the longitudinal axis of 
the missile is maintained at an angle to the trajec¬ 
tory, the missile will ultimately reach equilibrium as 
a glider traveling at constant speed (the terminal 
speed for this orientation) in a straight path which is 
inclined at an angle to the vertical. If propelled, the 
path may be inclined upward. Even if the controls 
are maintained in the neutral position, the missile 
will finally approach an equilibrium state of unac¬ 
celerated fall vertically downward at its terminal 
speed. For an ordinary bomb or other missile of high 


wing loading the time required corresponds to a fall 
through a very great height. 

The lifting missiles are intended to follow an ap¬ 
proximately straight flight path, which for powered 
missiles may be horizontal or inclined upward. We 
shall consider first an unpowered missile, i.e., a glider. 

The only forces acting on a glider in flight are the 
force of gravity and the resultant air force. It is cus¬ 
tomary to consider the resultant air force in terms of 
its components perpendicular and parallel to the di¬ 
rection of motion, the lift and the drag. In equilib¬ 
rium gliding flight the resultant of lift L and drag D 
must balance the weight and hence must act in the 
vertical direction. The flight path is therefore inclined 
downwards at an angle 0 such that tan 0 = D/L. 
The'resultant force R is usually expressed in terms 
of the dimensionless coefficient Cr defined by 
R = CrA^pV 2 , where A is the wing area, p the air 
density, and V the flight speed. Since at equilibrium 
the resultant equals the weight W, the equilibrium 
speed of the glider along its flight path is determined 
by the relation 

F = l/EJL 

y a P c R 

For an airplane the value of Cr increases from a 
small value at zero degrees or some small negative 
angle of attack, depending on the shape of the wing 
section, nearly linearly up to angles of the order of 
12 to 15 degrees, reaches a maximum value of the 
order of 1.2 to 1.4, and then slowly decreases. The 
ratio of lilt to drag reaches a rather sharp maximum 
value at some angle between 5 and 10 degrees and 
then decreases rapidly. The slowest equilibrium glid¬ 
ing speed corresponds to the maximum value of C R 
near the stalling angle, and in this region of angle of 
attack the glide angle changes rather rapidly but 
with only small changes in equilibrium speed. As the 
angle is reduced, the glide angle becomes flatter, and 
the equilibrium speed increases. On passing the angle 
of maximum L /D the glide path again becomes 
steeper, and the equilibrium speed increases still 
more. The steepest path is the vertical dive, in which 
the equilibrium speed reaches its maximum value, 
the terminal speed at which the weight is balanced 
by the drag. 

The preceding description applies solely to steady- 
state conditions, i.e., those which occur after the 
lapse of a sufficient interval of time. It is important 
to observe that an increase in angle of attack at an¬ 
gles beyond that of maximum L/D at first flattens 





MANEUVERABILITY OF VEHICLE AND OTHER AERODYNAMIC PROBLEMS 


241 


the angle of glide or even causes a temporary ascent, 
and the steeper glide angle is obtained only after the 
lapse of sufficient time. 

Let us suppose that while the airplane is gliding 
steadily at some angle of attack, the angle is sud¬ 
denly changed to a new value. The lift at angles be¬ 
low the angle of attack for maximum L/D is changed 
much more than the drag, and the principal effect 
will be to produce an unbalanced force nearly perpen¬ 
dicular to the direction of motion, which will curve 
the flight path. Ultimately of course the unbalanced 
drag force will modify the speed but this process re¬ 
quires some time. Suppose the lift coefficient corre¬ 
sponding to the steady, straight flight at the instan¬ 
taneous altitude and speed is C l° and the actual lift 
coefficient is Cl. The unbalanced force is then 
(Cl — C L°)A%pV 2 , and hence the lateral accelera¬ 
tion will be [(Cl — C L °)A%pV 2 ]/ ( W/g ). The radius of 
curvature of the path will be W/(Cl — C L °)Ag^p. 
This corresponds to the value Z\W / A given earlier, 
if Cl — Cl° is taken as 0.84. For a given wing loading 
and air density, the radius of curvature of the flight 
path depends on the difference between the actual 
value of the lift coefficient and the value which would 
be necessary in steady flight. 

The minimum radius is obviously obtained with 
Cl° = 0 and Cl equal to the maximum lift coefficient, 
i.e., with a nonlifting missile. The use of a lifting mis¬ 
sile increases the minimum radius of curvature and 
hence gives lower maneuverability. 

For angles of attack beyond the angle of maximum 
L/D, the effect of the drag may predominate, and the 
speed may be reduced so rapidly that the lift force 
is not increased, although the lift coefficient is greater. 

The foregoing discussion applies to control of the 
path of the missile in a vertical plane. The lifting 
missile usually has larger surfaces, i.e., wings for sup¬ 
port and control in the vertical plane. For right-left 
steering the lifting missile is usually designed to use 
the airplane method, i.e., banking or rolling the mis¬ 
sile to obtain a component of the force on the large 
wing surfaces in the desired direction. The turns 
which can be produced without banking are of very 
large radius. An airplane with dihedral angle will au¬ 
tomatically bank if the rudder is turned and ailerons 
are not moved, and will turn to right or left if the 
airplane is rolled either by deflecting the ailerons or 
by turning the rudder. 

Turns of a glider are descending spirals. If the in¬ 
clination of the spiral flight path to the horizontal 
is 6 and the radius of the spiral is r, the radius of 


curvature R of the flight path is r/cos 6. Call the 
angle of bank <£ and the lift L. The available force is 
then L sin (f> and hence 

, . W V 2 

L sin =- — 

9 R 

V being the flight speed, W the weight of the air¬ 
craft, and g the acceleration of gravity. Setting 
L = Cl\pAV 2 and solving for R 


p A C L g sin </> 

The radius of turn can be decreased by increasing 
Cl as the glider is banked. This is the method com¬ 
monly used in aircraft to make a tight turn. If C L 
is zero, i.e., a nonlifting missile, banking does not give 
rise to a turn, R being infinite. If Cl is not zero, i.e., 
a lifting missile, banking gives rise to a turn even if 
Cl is not increased by action of the longitudinal con¬ 
trol. This is one of the essential differences between 
the aerodynamic properties of lifting and nonlifting 
missiles. The right-left steering of lifting missiles by 
banking will be discussed further in Section 12.4. 

It is quite possible to design a lifting missile which 
could turn without banking. A large vertical surface 
would be needed, and various practical difficulties 
would arise. 

Nonlifting missiles will probably have less diffi¬ 
culty with compressibility effects at high speeds. The 
chief difficulty with the lifting missile arises from the 
fact that the missile is unsymmetrical about the plane 
of the wings, and the orientation is maintained by re¬ 
actions on wings and tail whose moments about the 
center of gravity are equal and opposite. Compres¬ 
sibility effects first appear on the wings, usually pro¬ 
ducing diving moments and large changes in atti¬ 
tude. The nonlifting missile, on the other hand, is 
symmetrical, and both body and fin moments are 
zero, except when control is applied. Thus the trim 
position of the nonlifting missile is not likely to 
change with speed in the absence of control. The con¬ 
trol will undoubtedly be modified by compressibility 
effects. 

At the present time aerodynamic data at super¬ 
sonic speeds are quite limited, consisting mainty of 
information on the drag of projectiles obtained in 
firing tests. Adequate supersonic facilities are now 
being provided, and within the next year or two a 
great deal of the necessary basic research should be 
completed. 




242 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


12 3 6 Powered Missiles 

The control of powered nonlifting missiles requires 
little further discussion. When no control is applied, 
the propulsive force acts in the direction of motion. 
Variation of the propulsive force changes the speed 
but not the direction of motion. When control is ap¬ 
plied, the orientation may change, in which case the 
propulsive force has a component at right angles to 
the direction of motion of the center of gravity, in¬ 
creasing the lateral force available for control. 
This effect is usually not large. 

In the case of powered lifting missiles, variation of 
the propulsive force constitutes a second and inde¬ 
pendent method of control in the vertical plane. Un¬ 
der equilibrium conditions the speed is approximately 
independent of the value of the propulsive force, sat¬ 
isfying the equation 

y _ 2 W cos 6 
~ ApCl 

For horizontal flight the propulsive thrust T is equal 
to the drag D. If T is greater than D, the missile 
climbs at angle 0 such that T = D + W sin 0. The 
angle 0 may be negative, corresponding to descent if 
T is less than D. For a more complete discussion, 
reference should be made to one of the many books 
on the subject of airplane performance. 

12 3 7 Homing Missiles 

In homing missiles, the apparatus contained within 
the missile locates the target with reference to some 
axis fixed in the missile. The information so obtained 
is not adequate, unless this reference axis is always 
in a known relation to the flight path of the center of 
gravity of the missile. The simplest solution is ob¬ 
viously to make the reference axis coincide with the 
flight path, i.e., to have the intelligence device “look” 
along the flight path. The installation is especially 
simple if the application of control does not change 
the orientation of the missile with respect to its flight 
path. The securing of this result is one of the aero¬ 
dynamic problems peculiar to homing missiles. The 
solution used in the Pelican and Bat projects was to 
change the lift of the wing by trailing edge flaps and 
so to locate the center of gravity of the missile and a 
fixed tail structure that the downwash effects on the 
tail produced moments counterbalancing the mo¬ 
ments produced by the flap deflection. The Roc proj¬ 
ect adopted the same solution. 


In the flight of an unpowered homing missile in 
still air against a stationary target, the flight path 
is approximately rectilinear. The initial speed is usu¬ 
ally much less than the terminal speed in the straight 
glide path, and hence the speed increases. To main¬ 
tain the rectilinear flight, the lift must be main¬ 
tained constant and equal to W cos 0. As the speed 
increases, the lift coefficient must be decreased by 
the action of the controls. 

It appears desirable to operate unpowered homing 
missiles within the range of lift coefficients lying be¬ 
tween zero and that for maximum lift-drag ratio. As 
is well known, the equilibrium flight path first flat¬ 
tens and then steepens as the lift coefficient is in¬ 
creased. There is accordingly a reversal of control as 
regards the final effect when past the maximum lift- 
drag ratio. Missiles are not usually in equilibrium on 
their flight path, and computation shows that the 
first effect of control is always that of the change in 
lift. If, however, the lift-drag ratio at the maximum 
lift coefficient permitted by the controls corresponds 
to a slope of path steeper than the actual path slope 
to the target, the control proceeds to its limit and 
stays there. The intelligence calls for a higher lift. 
A higher lift coefficient is obtained as the control 
moves toward its limit, but a higher drag coefficient 
also results. The drag reduces the speed so that the 
actual lift does not increase very fast, and the flight 
path is corrected very slowly, if at all. It is probably 
desirable that powered homing missiles should also 
be operated in the region between zero and maximum 
lift-drag ratio. 

If the homing missile is powered, various combina¬ 
tions of control are possible. For example, the pro¬ 
pulsive thrust might be controlled by a speed-sensi¬ 
tive device to maintain a constant speed in the later 
stages of the flight. 

12 3 8 Stability Problems 

In theory a homing missile might receive sufficient 
control information to give stable flight without in¬ 
herent stability of the missile in the absence of con¬ 
trol. In practice, a missile must have satisfactory 
stability to maintain its flight in periods of inten¬ 
tional or unintentional absence of control informa¬ 
tion. 

It is not practical to review the many aspects of the 
stability of missiles. The disturbed motions of a 
stable airplane-like missile in the absence of homing 
control consists of various combinations of a rapid, 





COUPLING BETWEEN CONTROLS 


243 


heavily damped, longitudinal motion (in which the 
angular pitching motion is most prominent), a slow 
oscillation (the previously mentioned phugoid oscil¬ 
lation, in which the missile rises and falls with the 
speed decreasing and increasing), a rapidly damped 
rolling and yawing oscillation, and a slow spiral mo¬ 
tion. 

Under certain conditions the missile may pass from 
steady rectilinear flight to a steady spin. In a true 
spin, as distinguished from the spiral motion referred 
to above, the wing surface is stalled, i.e., it meets the 
air at a large angle. The spinning motion is, however, 
a steady motion Avith the following balance of mo¬ 
ments and forces: 

1. The stalled wing rotates of itself at such a speed 
that the rolling moment is zero. (A wing at normal 
angles offers resistance to rolling; a stalled wing is in 
unstable equilibrium at zero rate of roll.) 

2. The aerodynamic pitching moment that tends 
to reduce the angle of attack is balanced by the cen¬ 
trifugal pitching moment that tends to increase the 
angle of attack. 

3. The aerodynamic yawing moment is balanced 
by the centrifugal yawing moment. The spinning 
characteristics are greatly affected by the angle of 
yaw at which this balance occurs. 

4. The airplane descends at such a rate that the 
vertical component of the air force equals the weight. 

5. The horizontal component of the air force gives 
the requisite centripetal acceleration of the center of 
gravity towards the spin axis. 

A missile can be made difficult to spin by a suitable 
disposition of tail surfaces to give large aerodynamic 
pitching moments and a favorable antispin equilib¬ 
rium angle of yaw. 

When a homing device is applied to an airplane¬ 
like missile, the disturbed motions take on a some¬ 
what different character. For example, if only the 
longitudinal stability is considered, the slow phugoid 
oscillation disappears and is replaced by a damping 
of any disturbance of the velocity along the flight 
path; the period and damping of the fast angular 
pitching oscillation are controlled in part by the 
static longitudinal stability of the missile and in part 
by the power of the control surface and the lag of the 
servomechanism; and there arises a second angular 
pitching oscillation of longer period, which may in 
some cases degenerate into two aperiodic motions. 
As the missile approaches the target, the frequencies 
and damping constants change somewhat, and an 
additional mode appears. 


In a similar manner the slow spiral motion asso¬ 
ciated with the lateral stability of the free-flying mis¬ 
sile disappears. There are short- and long-period, 
combined rolling and yawing oscillations, where 
damping may be positive, zero, or negative, and the 
long-period motion may degenerate into aperiodic 
motions. 

Some aspects of the stability of homing missiles 
will be discussed briefly in Section 12.5. 

124 COUPLING BETWEEN CONTROLS 
12,41 Interdependence of Controls 

A missile in flight has six degrees of freedom, but 
if it follows conventional aircraft design it has only 
three controls—unless it is propelled, in which case 
it has also a throttle or equivalent thrust control. 
The three controls are usually movable surfaces to 
control the moments about three mutually perpen¬ 
dicular axes, and there is no direct control of the 
linear motions. Unfortunately the three controls do 
not give independent effects. Thus the rudder pro¬ 
duces a small rolling moment as well as a yawing 
moment which turns the aircraft to right or left. 
When the missile is yawed, a much larger indirect 
rolling moment results, arising from the aerodynamic 
reactions. The ailerons produce yawing moments as 
well as rolling moments, and the yawing moments 
may be either favorable or “adverse,” i.e., the re¬ 
sultant yaw may produce a rolling moment in a direc¬ 
tion opposite to the desired rolling moment. In ex¬ 
treme cases the ailerons may turn the aircraft to 
right or left without rolling it, or the rudder may roll 
the aircraft with little yaw. The effects may change 
sign for the same aircraft at different speeds, i.e., at 
different angles of attack or at different lift coeffi¬ 
cients. The interaction between the rolling or yaw¬ 
ing motion and the pitching motion is fortunately 
very small. 

In the case of nonlifting missiles, operation of the 
left-right and up-down controls simultaneously pro¬ 
duces rolling moments, and hence it has been found 
that a nonlifting missile must have ailerons if rolling 
motions are to be avoided. 

0 

12,4,2 Intelligence Coordinates and Controls 

The intelligence devices generally available for use 
in homing aero-missiles give information on the bear¬ 
ing of the target to the right or left and up or down 




244 


SOMk ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


from some reference axis, i.e., an essentially two- 
dimensional presentation. In the case of radar de¬ 
vices, range information can also be obtained if de¬ 
sired. The missile, however, has three controls if 
unpropelled and at least four if propelled. There is 
evidently a problem in connecting the two-dimen¬ 
sional output of the intelligence device with the con¬ 
trol system of the missile. 

In the case of a nonlifting missile, the obvious 
method of connecting the controls is to connect the 
two channels of the intelligence device to two con¬ 
trol surfaces producing moments about two mutu¬ 
ally perpendicular lateral axes, which are aligned with 
the intelligence device. While the two control sur¬ 
faces may be designed to give independent action 
about the lateral axes, the application of control 
about both axes simultaneously will produce rolling- 
moments about the longitudinal axis. If the aero¬ 
dynamic design were such that the longitudinal axis 
remained in the direction of motion, the roll would 
not be objectionable, unless the rate was so high that 
the lag in the control system introduced phase er¬ 
rors. However, most missiles of the zero lift type 
change orientation as the controls are applied. The 
roll takes place about the axis of the missile and intro¬ 
duces incorrect error signals in the intelligence de¬ 
vice. A better solution is to stabilize the missile in 
roll by means of ailerons controlled from a suitable 
gyro system. Thus in the simplest method, the roll 
control is governed by a gyro, and the other two by 
the channels of the intelligence device. 

There have been many ingenious suggestions for 
connecting the controls of a nonlifting missile to per¬ 
mit continuous rotation of the missile. It does not 
seem profitable to discuss them here. For remotely 
controlled as distinguished from homing missiles, 
such devices become computers, often of complex de¬ 
sign, for transforming from fixed to rotating axes, 
and metering the degree of control to be given by 
each of the two control surfaces. 

In the case of a lifting missile of the airplane type, 
turns are accomplished by banking the airplane, and 
it is therefore not desired to prevent roll. It is neces¬ 
sary either to reduce the number of control surfaces 
from three to two, or to couple two of the control 
surfaces together to be controlled from a single intel¬ 
ligence channel. If the missile is powered, the throttle 
or equivalent power plant control may be arranged 
to be controlled by airspeed, altitude, or some quan¬ 
tity associated with engine performance as for ex¬ 
ample, mixture ratio, peak pressure, temperature, 


etc. Combinations may be used but it has not been 
customary to use the throttle for up-down control of 
the flight path. 

The more common two-control airplanes use the 
elevator-rudder or the elevator-aileron combinations. 
Both methods have been used in the design of mis¬ 
siles. With proper design reasonably satisfactory 
turns can be made either with ailerons or with rud¬ 
der. For homing missiles, aileron control is believed 
to give somewhat smaller errors, since out-of-trim 
rolling moments can be balanced without introducing 
yaw. In either case, the use of two controls alone 
makes possible independent connections of the two 
intelligence channels. 

It has already been pointed out that a gyro is re¬ 
quired for trimming the missile in roll. It is desirable 
that this gyro also limit the maximum angle of roll; 
otherwise the missile is likely to turn over on its back 
when a control signal is continued for some time. 

In automatic pilots for aircraft, the rudder and 
aileron are often controlled together for turns. Such 
a coordination is possible for one flight condition or, 
at most, a narrow range of conditions. For missiles 
intended to operate over a considerable speed range, 
the rudder displacement for a given aileron displace¬ 
ment varies both with airspeed and angle of attack, 
and it has not seemed necessary to attempt the de¬ 
sign of the necessarily complicated control. By suit¬ 
able aerodynamic design of a two-control missile, the 
sideslip during turns can be made reasonably small. 

12 4 3 Effect of Roll on Error Signals 

The lifting missile of the airplane type banks dur¬ 
ing a turn. The reference axes of the intelligence de¬ 
vice are fixed with reference to the missile and hence 
rotate with the missile. The axis of rotation of the 
missile does not usually coincide with the direction 
of motion of the center of gravity of the missile. 
Hence the intelligence device no longer measures the 
up-down and right-left errors with respect to axes 
fixed with respect to the vertical. 

It has been previously pointed out that an airplane 
is made to turn in its tightest circle by banking and 
then pulling back on the stick to increase the lift co-/ 
efficient to its maximum value. The coupling between 
the up-down and right-left controls produced by 
banking automatically gives an error signal to the 
elevator or elevons in a direction to increase the lift 
coefficient and thus accelerates the turn. At the same 
time the increased lift bends the trajectory of the 



SYSTEM STABILITY OR HUNTING PROBLEMS 


245 


center of gravity upward. The exact behavior of the 
system depends on the characteristics of the intel¬ 
ligence device, servomechanism, and missile. 

The error angles are readily computed for idealized 
motions. If the missile rotates about the direction of 
motion of the center of gravity, the effects amount 
simply to rotation of the axes of reference of the error 
signals. If the error angles referred to the original 
axes are 8 e in elevation and 8 a in azimuth, and if re¬ 
ferred to axes rotated through an angle 0 they are 
8/ and 8 a f , the relations are 

8 a ' = 8 a cos <t> — 8 e sin 0 

8 J = 8 a sin 0 + 8 e cos 0 

If, however, the rotation occurs about an axis mak¬ 
ing an angle 8 eo with the direction of motion of the 
center of gravity 

8 a ' = 8a cos 0 ( 8 e 8 cq ) sin 0 

8 e ' = 8 a sin 0 + (8 e — 8 eo ) cos 0 

In the general case the dynamics of a rigid body 
with six degrees of freedom and the aerodynamic 
characteristics of the control must be considered. 
Thus the axis of torque varies with the displacement 
of the control surfaces. The body rotates not about 
the axis of torque, unless this axis is also a principal 
axis of inertia, but about an axis intermediate be¬ 
tween the axis of torque and a principal axis. Thus 
the instantaneous axis of rotation travels in the body, 
and an integration process is necessary. Because of 
these complications it is not very practical to correct 
for these effects by inserting a computing device be¬ 
tween the intelligence and the controls. One of the 
important consequences of the foregoing effects is 
that when there is an elevation error angle and the 
true azimuth angle is zero, an azimuth error is passed 
on to the controls equal to — (8 e — 8 eo ) sin 0. To 
illustrate, suppose that the target is high, so that 
8 e — 8 eo is positive, and that the airplane rolls as for 
a right turn (positive 0). A negative azimuth error is 
then given to the controls, which tends to oppose the 
right turn. If, however, 8 e — 5 eo were negative, i.e., 
true elevation error is zero but axis of roll above 
flight path or target is low, the azimuth error given 
to the controls would assist the turn. In the first case 
there would be a damping effect on oscillations; in 
the second, a destabilizing effect which would pro¬ 
mote hunting. In the usual lifting missile the effect 
is on the average a destabilizing one. 


12 4 4 Effect of Angle of Attack 
or Yaw on Error Signals 

In the nonlifting missile, the application of con¬ 
trols pitches or yaws the longitudinal axis so that the 
intelligence no longer looks along the flight path. 
Using the same notation as before, the effect is given 
by the relation 

8 e = 8 e — 8 CQ 

The same relation applies to the elevation control of 
a lifting missile. It is in theory possible to intercon¬ 
nect the controls with the intelligence device to move 
the reference axes of the intelligence device to com¬ 
pensate for this effect. If it is undercompensated 
there will be a stabilizing action against oscillations; 
if it is overcompensated a destabilizing action. 

12.4.5 Methods of Reducing Coupling 

between Controls 

The methods of reducing both the interdependence 
of the controls and the undesired effects of the angu¬ 
lar motions on the error signals are still in the early 
development stages, and much additional aerody¬ 
namic research is required. A homing missile can be 
made to work with all these effects present in some 
degree. It would seem fruitful, however, to attempt 
to make the normal flight path coincide with the 
longitudinal principal axis of inertia for all control 
positions, and to do further research on reducing 
interactions between controls. 

12.5 SYSTEM STABILITY 
OR HUNTING PROBLEMS 

12 51 Illustration of Hunting 

Perhaps the most difficult of the special problems 
associated with the design of homing missiles is that 
of securing a satisfactory stability of the complete 
system, since the overall stability depends on the 
characteristics of the missile, intelligence, and servo¬ 
mechanism and especially on their interrelations. To 
illustrate the problem there are shown in Figures 2, 
3, and 4 the observed motions of three homing mis¬ 
siles. In Figure 2 it is seen that the missile rolls and 
yaws in increasing oscillations of about a 10-second 
period until the missile strikes the ground. In extreme 
cases, such a missile may turn completely over on its 
back, lose track of the target and descend in a steep 



246 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


APPARENT ANGULAR HORIZONTAL MOTION OF BEACON 




Figure 2. Unstable oscillation of homing missile. 


spiral. In Figure 3, the missile pitches in a steady os¬ 
cillation of nearly constant amplitude and a period 
from 5 to 7 seconds. Figure 4 shows a rolling and 
yawing oscillation which is damped as the flight 
progresses. The period of the main oscillation is ap¬ 
proximately 10 seconds but there is superposed a 
rapid roll oscillation of about 4 degrees amplitude 
and 134 -second period. 

The steady hunting motions illustrated in Figures 
3 and 4 are objectionable only as they affect the path 
of the center of gravity of the missile. The unstable 
motion of Figure 2 usually, but not always, results in 
complete failure of the missile and a wide miss. 

The effect of a given hunting motion on the tra¬ 
jectory may be estimated as follows. The error angle 


is equal to the ratio of the velocity dy/dt of the mis¬ 
sile transverse to the flight path to the velocity V of 
the missile, provided the missile looks along its flight 
path. If this angle oscillates sinusoidally with maxi¬ 
mum amplitude do, the motion may be represented by 


dy 2t rt 

Tt = Ve „co ST - 


where T is the period. The lateral displacement y is 
then given by the formula 


V = 2/o + 


VOpT 

2t 



where yo is the mean value of y. The maximum excur¬ 
sion from the trajectory is then VdoTfor. For 





































































SYSTEM STABILITY OR HUNTING PROBLEMS 


247 


V = 400 ft per sec, 0 O = 3 degrees = 1/20 radian, 
and T = 2tt seconds, the flight path oscillates ap¬ 
proximately 20 ft. On the other hand, if T = 1 sec¬ 
ond, the flight path oscillation is only 3 ft. 

These simple computations underestimate the er¬ 
ror since the actual missile does not look accurately 
along the flight path under dynamic conditions. This 
effect is, however, greatest for the short-period oscilla¬ 
tions where the errors are small. 


12,5,2 Factors Leading to Instability 

The presence of error signals from a homing de¬ 
vice would seem at first sight to remove most of the 
sources of instability and to supplement the aero¬ 
dynamic damping forces which are present in any 
normal design. The element which introduces diffi¬ 
culty is the unavoidable lag between the presence of 
an error angle and the application of the corrective 


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248 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


forces and moments to the missile. The lag may arise 
in the mechanical and electrical parts of the intel¬ 
ligence device, control system, or servomotor, or it 
may be introduced by the motion of the missile re¬ 
sponding to the controls in such a manner as to intro¬ 
duce errors in the measurement of the error angle by 
the intelligence device. 

The physical effect of the lag in producing hunting- 
may be appreciated by considering the missile dis¬ 
placed from the correct flight path toward the target 
but being returned to it by the action of the controls. 
As the error angle approaches zero, the control sur¬ 
face deflection is approaching zero and should reach 
zero at the same instant as the flight path intersects 
the target. If the control is of the on-off type, re¬ 
versal should occur at the instant the flight path in¬ 
tersects the target. If this could be effected, the aero¬ 
dynamic damping forces would absorb some of the 
kinetic energy present, and the next crossings would 
be made with successively decreasing speed until the 
missile reached equilibrium on a steady trajectory. 
Because of the inevitable lag present, the control sur¬ 
face remains deflected in a direction to increase the 
error angle of the missile for a short time and thus 
feeds energy into the oscillatory motion. Equilibrium 
is reached at an amplitude of oscillation such that 
the amount of energy fed in is equal to that absorbed 
by the aerodynamic damping. It is possible that the 
energy fed in does not equal the energy absorbed for 
any amplitude, in which case the oscillation builds up 
until the missile capsizes. 

The amount of energy fed in is directly propor¬ 
tional to the time lag, and to the overall sensitivity, 
i.e., to the magnitude of the correcting forces and 
moments produced by a given error angle. This over¬ 
all sensitivity is dependent on such factors as the 
area of the control surfaces, the speed, altitude, and 
attitude of the missile, the deflection or rate of de¬ 
flection of the control surface for a given electrical 
input to the servomechanism, and the sensitivity of 
the intelligence device. 

The aerodynamic damping is a function of altitude 
and airspeed as well as of relative areas of various 
parts of the missile and of the amplitude of oscilla¬ 
tion. 

The time lag in the intelligence device is usually 
determined by the amount of integration or smooth¬ 
ing found desirable to give a reasonably sioady out¬ 
put signal from the intelligence device. For the radar¬ 
homing systems used so far, the time lag has been 
within the range from 0.05 to 0.25 second. 


The time lag of the usual electromechanical servo 
system is of the order of 0.10 to 0.15 second, but con¬ 
siderably smaller values can be obtained by special 
design. 

Either a time lag or a time advance may be intro¬ 
duced by the relations between the motions of the 
missile, the axis of the intelligence device, and the 
direction of motion of the center of gravity of the 
missile, as discussed in Section 12.4.3. In the usual 
design, the axis of roll lies between the longitudinal 
axis of the missile and the direction of motion of its 
center of gravity, and a time lag is introduced. For 
example, if a disturbance rolls the missile for a right 
turn, a false azimuth error is indicated in a direction 
to aid the turn and amplify the effect of the dis¬ 
turbance. From the point of view of lag, if the missile 
is off course to the left and is being corrected to the 
right, the centering or reversal of control is delayed 
until the missile moves off course to the right a suffi¬ 
cient amount to bring the rotated reference axis on 
the target. 

The use of automatic lead computers to reduce 
errors associated with wind and moving targets often 
leads to a difficult stability problem, as discussed 
later in this chapter. 

12 5 3 Antihunt Devices 

Since hunting indicates a phase lag between the 
error angle and the application of corrective forces 
and moments, the remedy is obviously the introduc¬ 
tion b}^ some means of a phase advance which is 
greater than the lag. As the missile comes on course, 
the control must be reduced or removed before the 
error is actually zero. The most common method of 
accomplishing this result is to incorporate a rate 
term in the control, i.e., to make the position or speed 
of the control dependent in part on the error and in 
part on the rate of change of the error. 

A compensating or antihunt circuit can often be 
introduced in the amplifier or output circuit of the 
intelligence device itself. It usually takes the form of 
a suitable condenser-resistance network which may 
be considered either as a phase-advancing device for 
sinusoidal signals or as introducing a rate component 
for arbitrary signal variation. Difficulties often arise 
with such a circuit if the input signals are “rough” 
(fluctuating in magnitude for a given error angle), 
as, for example, is the case for radar reflections from 
many types of targets. Because of the antihunt cir¬ 
cuit, the higher frequency “roughness” is amplified 



SYSTEM STABILITY OR HUNTING PROBLEMS 


249 


much more than the slower variation as the missile 
comes on course. A further characteristic of this anti¬ 
hunt circuit is that good performance is obtained 
only with considerable attenuation of the input sig¬ 
nal, so that additional amplification is required. 

The rate term may be introduced by a gyroscopic 
turn indicator which measures the angular rate at 
which the missile comes on course. This method has 
been found completely effective in obtaining system 
stability. It has the disadvantage that winds or mov¬ 
ing targets require a steady rate of turn, which intro¬ 
duces a control signal. The presence of this signal, in 
effect, limits the maximum rate of turn and thus in¬ 
creases the errors associated with winds and moving 
targets. 

12.5.4 Time to Come on Course 

One of the important characteristics of the control 
system is the time required for a missile released with 
an initial error angle to come on its proper course. 
The time can be controlled to some extent by the 
overall sensitivity of the control system. It has al¬ 
ready been pointed out that increasing the overall 
sensitivity increases the amount of energy fed into 
the system when a time lag is present. Even if the 
system is stable, the missile may come on course with 
a damped oscillatory motion, and the damping is 
small if the sensitivity is too great. A return without 
oscillation can be secured with suitable values of the 
control parameters. Both experience and theory show 
that it is necessary to make a compromise between 
the stability characteristics and the time to come on 
course. In most cases it is necessary to make the time 
to come on course about 10 seconds or more. 


12 5 5 Effect of Wind Gusts 

Wind gusts introduce disturbances which excite 
the natural modes of motion of the complete system. 
The effects persist for times which depend on the 
periods or time constants of the natural modes, and 
the disturbed motion at any instant depends on the 
history of disturbances over comparable past periods 
of time. Exclusive of the effects of changes in speed 
along the trajectory for which the time constant is 
very large but which in themselves do not cause the 
missile to miss the target, there usually exist modes 
for which the time constant is of the order of 10 
seconds. Hence gusts in the latter part of the flight 


path introduce errors which cannot be wholly cor¬ 
rected in the time available. 

Experience has shown decreased dispersion in tests 
of homing missiles over water as compared with tests 
over land, an effect which is probably to be attributed 
to the decreased gustiness over water. In one instance 
over land, a severe disturbance was noted from the 
sharp boundary between a heated layer of air at the 
ground and a cold air mass aloft. While the designer 
of a missile can vary the frequencies of the natural 
modes of motion through small limits, it is practically 
impossible to change their order of magnitude. For¬ 
tunately the errors in the trajectory decrease as the 
rectilinear flight speed increases. 

12,5,6 Methods of Analysis 

The problem of the design of stable systems has 
been approached by several methods of mathemati¬ 
cal analysis and by experiments on mechanical, elec¬ 
tromechanical, or electrical models. If sufficient in¬ 
formation is available from tests of the component 
parts, the choice of control parameters to assure sys¬ 
tem stability can be made by any of these methods 
of analysis. The purely mathematical methods re¬ 
quire more or less idealized representation of the per¬ 
formance of the component parts of the system and 
are most useful for systems whose performance is de¬ 
scribed by linear differential equations. The various 
methods using models have the advantage that some 
of the actual components can be incorporated as part 
of the model, so that nonlinear control mechanisms 
and on-off or step controls whose mathematical anal¬ 
ysis is often difficult can readily be investigated. 

The control system of a homing missile constitutes 
a closed-cycle control system or servomechanism. 
The chief difference from ordinary servomechanisms 
is that the parameters of the system are not constant 
but vary through considerable limits during the flight 
of a single missile. More specifically, the parameters 
associated with aerodynamic forces and moments are 
functions of the airspeed, air density, and attitude 
of the missile and its control surfaces. If these varia¬ 
tions are taken into account, the usual methods of 
analysis of servomechanisms may be applied. These 
involve a study of the response of the system to cer¬ 
tain standardized conditions, the two most useful 
conditions translated in terms of a homing missile 
being a sinusoidal displacement of the target at vari¬ 
ous frequencies or a sudden permanent displacement 
of the target. If the system is a linear one, its per- 



250 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


formance is completely known when the performance 
under either of the above conditions is known. 

Readers who are interested in the details of the 
mathematical procedures are referred to reference 2. 
This paper describes the application of the sinusoidal 
analysis to the design of linear servomechanisms 
with continuous control. An example of analysis using 
suddenly applied disturbances may be found in ref¬ 
erence 3. The same laboratory has developed this 
method of studying the control of certain types of 
missiles. 

The preceding methods of analysis lend themselves 
to a complete study of the performance of the sys¬ 
tem, including not only the question of whether the 
system is stable or not but all questions of the magni¬ 
tude of the errors. For determination of stability, one 
may apply the method of small oscillations commonly 
used in airplane stability problems. This method as 
applied to an airplane without automatic pilot is de¬ 
scribed in great detail in reference la. The method, 
in effect, determines the period and damping of the 
natural modes of oscillation and the damping of the 
natural aperiodic motions. If the damping constant 
turns out to be negative, the corresponding mode is 
unstable. 

The mathematical methods are in practice limited 
to systems described by linear differential equations 
and find some difficulty in the treatment of actual 
servo systems with friction, “dead” regions, and time 
lag. For this reason models have been found useful. 
These range in complexity from systems representing 
a single degree of freedom of the missile to more com¬ 
plete flight tables, which include the three angular 
components of the motion or the “phantasmagoria” 
which simulate one or two components of the linear 



Figure 5. Diagram of launching variables determin¬ 
ing pursuit curve errors. 


motion. As an example, methods of automatic roll 
stabilization may be investigated with the actual 
gyro and servo elements by using a mechanical sys¬ 
tem (damped pendulum or damped rotor) to simu¬ 
late the inertia and damping of the missile about its 
roll axis. A system of this type and a more compli¬ 
cated electromechanical model of the longitudinal 
motion were used in the study of the Pelican and Bat 
control systems. In addition, the Servomechanisms 
Laboratory, Massachusetts Institute of Technology, 
developed a flight table in which the pitching, yaw¬ 
ing, and rolling motions of the missile with proper 
cross coupling were fully simulated and on which an 
overall test could be made of the complete control 
system including the intelligence unit, gyroscopes, 
and servomechanism. This flight table was extremely 
useful in determining the best adjustments of the 
several parameters of the control system. 

Comparison of these models with actual flights of 
homing missiles shows that they reproduce the an¬ 
gular motions encountered in flight extremely well. 
Designers of homing missiles \fall find that the use of 
this method of studying system stability and overall 
performance will save much time in the development 
of a stable system and in adjusting it for best per¬ 
formance. 

12 6 PROBLEMS ARISING FROM 
WIND AND MOVING TARGETS 

Pursuit Curves 

If the target is moving or there is a natural wind, 
a flying missile which always heads toward the target 
follows a path known as a pursuit curve or homing 
curve, depending on whether the motion is referred 
to the air or to the ground. The radius of curvature 
of the pursuit curve becomes very small as the target 
is approached and becomes infinitely small for the 
idealized case of a point target. The maneuverability 
of the missile is limited, and hence the missile will not 
hit a moving point target or a stationary target in a 
cross wind. The magnitude of the miss depends on 
the speeds of the missile, wind, and target, on their 
relative directions, on the range, and on the maneu¬ 
verability of the missile as expressed by its minimum 
radius of curvature. 

This problem in idealized form has been studied by 
the Statistical Research Group, Division of War Re¬ 
search, Columbia University, under the direction of 
the Applied Mathematics Panel. For convenience, 




PROBLEMS ARISING FROM WIND AND MOVING TARGETS 


251 


the effect of wind and target motion are combined to 
give an apparent target motion of speed v. (See Fig¬ 
ure 5.) The azimuth fo of the launching position re¬ 
ferred to the direction of apparent motion of the tar¬ 
get, the launching radius r 0 , the ratio n of the appar¬ 
ent target speed v, (the resultant of wind effect and 
target motion) to the speed of the missile V, and the 
minimum turning radius p m of the missile are the 
quantities determining the miss for the case of the 
point target. Constant velocities of target and mis¬ 


sile, direction of launching toward the target, and 
missile continuing along an osculating circle when 
the minimum radius of curvature is reached are 
assumed. 

The computed miss M is shown in Figure 6 in the 
form of a plot of M / p m vs </>o forn = 0.1, 0.2, 0.3, and 
0.4 and r 0 = 6 p m , 10p m , and 14 p m . The essential points 
to notice are: (1) that the maximum miss varies but 
little with launching radius; (2) that the maximum 
miss is approximately equal to }^p w n 2 ; and (3) that 



Figure 6. Pursuit curve errors due to effects of wind, target motion, and limited maneuverability of missile. 







252 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


the miss becomes very small for <f> 0 between 140 and 
180 degrees. The only factors within the control of 
the designer of the missile are the speed of the missile 
and its minimum turning radius. The speed ratio 
should be as small as possible and the minimum 
turning radius as small as possible to minimize pur¬ 
suit-curve errors. The user of the missile may reduce 
the error by making his attack on a course such that 
<f> o is approximately 145 degrees. In the case of ship 
targets this course gives a good compromise between 
the magnitude of the error and the projected length 
of the ship, which determines the permissible error 
that still yields a hit. 

12 6 2 Computers 

The essential feature of the pursuit curve is that 
the missile moves along a path which makes an angle 
tan -l (w sin 0 O )/(1 -f n cos 0 O ) with the line joining 
the missile and target. If the axis of the intelligence 
device were rotated with respect to the longitudinal 
axis of the missile by this amount in the opposite 
direction to the apparent target motion, the path 
with respect to axes fixed in the target would be 
straight, and the missile would travel on a collision 
course. If the user of the missile has the necessary 
data to compute n and <£ 0 and the means of offsetting 
the axis of the intelligence device by the computed 
amount, and if the variations of n and <j> 0 may be 
neglected, the pursuit-curve error may be eliminated. 

It is, in theory, possible to design a computer which 
automatically offsets the axis of the intelligence de¬ 
vice by the required amount. The simplest scheme, 
in theory, is to equip the intelligence device with a 
separate servo system which always keeps the axis 
of the device centered on the target. If the missile 
flies a collision course, the bearing of the target will 
remain constant; hence the angular velocity of the 
intelligence device will be zero. If the controls are so 
arranged that the missile is guided to make the angu¬ 
lar velocity zero, the course followed should be the 
desired collision course. In practice, a high precision 
is required, since the rate of change of bearing for a 
small offset is very small at long ranges. Further¬ 
more, the missile is subject to numerous disturbances 
of attitude, and it appears that an accurate, linear- 
integrating, angular-velocity meter would be re¬ 
quired. It is not known whether the scheme would 
work or not; it has not yet been tried on a missile. 

Another possibility which amounts to integrating 
the angular velocity is to determine the angular dis¬ 


placement of the axis of the missile after a certain 
time or distance by means of a free gyro which main¬ 
tains a fixed direction in space. A knowledge of the 
angular displacement and the range at the beginning 
and end of the displacement permits a computation 
of the desired offset. This computation also assumes 
a constant value of n throughout the flight. 

Still another possibility is a continuously integrat¬ 
ing angular-velocity meter which continuously sets in 
the appropriate offset. 

Study shows that the roughest sort of correction 
reduces the pursuit-curve errors, even though a true 
collision course is not obtained. Hence it is worth 
while to experiment with the simplest conceivable 
computers—for example, those dependent merely on 
the relative frequency and length of error signals in 
the two opposite directions. 

The chief bar to the general use of computers lies 
in their effect on system stability; in fact, the auto¬ 
matic devices described cannot be used because they 
produce excessive hunting. The computer essentially 
puts in a lead on a moving target. If the missile, as a 
result of a disturbance, rotates to increase the lead, 
the angular velocity developed acts to increase the 
lead still further and thus builds up an oscillation of 
increasing amplitude. The computer cannot distin¬ 
guish between angular velocity arising from disturb¬ 
ances and angular velocity caused by the missile fol¬ 
lowing a pursuit curve. Consequently an automatic 
lead computer is an inherently destabilizing device. 

The two methods of solving the stability problem 
are: (1) to feed in only a fraction of the correct lead 
angle, or (2) to introduce a time constant in the in¬ 
tegration process which is long compared to the nat¬ 
ural period of oscillation of the homing missile. In the 
first case, the permissible fraction is dependent on 
the stability reserve of the complete system of the 
intelligence device, servo control, and missile. In the 
second case, the time constant must be of the order 
of at least 20 or 30 seconds for missiles exhibiting 
natural periods of the order of 1 to 8 seconds. 

The computer problem is susceptible of mathe¬ 
matical analysis for any specific missile and control 
system, and it is not too difficult if all elements of the 
system are linear. Actual experience with computers 
on a homing missile is not yet available. 

12 7 BORE-SIGHT ERRORS 

If the axis of the intelligence device does not co¬ 
incide with the direction of motion of the center of 



BORE-SIGHT ERRORS 


253 


gravity of the missile, there arises an error which may 
be termed the bore-sight error. Let us confine our 
attention to the case where the intelligence device is 
not equipped with a computer or other means of 
introducing an offset angle, i.e., the axis is fixed with 
respect to the missile. The missile then flies at a fixed 
angle to the line joining the missile and the target 
and hence follows a logarithmic spiral whose equation 
in polar coordinates is p = p o e~ {e ~ 0o)/t&n e ’ where p 0 and 
6 0 are the coordinates of the release point and e is the 
bore-sight error. As the missile approaches the target, 
the radius of curvature decreases and finally becomes 
equal to and would be less than that which the mis¬ 
sile can follow. The radius of curvature of the path 
is equal to p /tan e or, if e is small, to p/e. Thus if p w 
is the minimum radius of turn of the missile, this 
value is reached at a value of p equal to p m e. This 
value of the range to the target at which the radius 
of turn becomes p m will be designated R m . If it is as¬ 
sumed that the missile continues to travel in a cir¬ 
cular path of radius p m , the miss is readily computed 
to be 3^Pm€ 2 . If the path were straight, the miss would 
be R m e = p m e 2 and the circular path produces a cor¬ 
rection 3^Pmt 2 , leaving a residual miss of ^p m e 2 . For 
pm = 10,000 ft and e = 1 degree, R m = 167 ft, and 
the miss is only 1.4 ft. The miss increases as the 
square of the bore-sight error. 

Bore-sight errors arise from many sources in addi¬ 
tion to the obvious one of inaccuracy in construction 
of the vehicle and of the mounting brackets of the 
intelligence device. Both mechanical and electrical 
imperfections of the intelligence device may produce 
bore-sight error. Thus, in a radar-homing device the 
electrical axis of the antenna system may not coin¬ 
cide with the geometrical axis because of unsymmet- 
rical distribution of dielectric or conducting material 
near the antenna. The output circuits may be un¬ 
balanced in such a manner that the output is not zero 
when the target is on the axis of the antenna. The 
servo control system may contain elements which are 
not balanced when the input signal is zero; for ex¬ 
ample, the pick-off of a rate gyroscope may be dis¬ 
placed from the correct zero position. Errors from 
these and similar causes may be controlled by careful 
inspection tests of the individual components, and 
an overall check of an installed system can readily 
be made. The reference axis of the vehicle can be 
oriented with respect to the target until the indicated 
output is zero, in which case the angular displace¬ 
ment of the target from the reference axis gives the 
bore-sight error. 


The most difficult problem is to determine that 
reference axis of the vehicle which lies in the direc¬ 
tion of motion of the center of gravity of the vehicle 
in free flight. In fact a vehicle designed like a conven¬ 
tional airplane travels at different attitudes for dif¬ 
ferent positions of the control surfaces. Particularly 
in the vertical plane, the angle of attack varies with 
the elevator setting. Thus, there is a variable bore- 
sight error which might amount to as much as 10 
degrees or more. 

A similar error may occur in the horizontal plane 
when rudder and elevator control are used, for in this 
case the rolling moment due to any lack of symmetry 
of the missile about its longitudinal axis must be bal¬ 
anced by application of the rudder to yaw the mis¬ 
sile, which then flies at an angle to its longitudinal 
axis. 

Two methods have been used to reduce bore-sight 
errors from aerodynamic causes. The first and most 
satisfactory is to design the missile so that the change 
in attitude with application of control is as small as 
possible. This requires aileron-elevator control rather 
than rudder-elevator control, so far as the horizontal 
motions are concerned. In the vertical plane the aero¬ 
dynamic design should be such that the lateral forces 
are modified without change in attitude. One method 
is to use flaps on the trailing edges of the wings to 
change the lift and to balance the pitching moments 
so produced by pitching moments from the tail aris¬ 
ing from changes in the down wash angle. This can be 
done by proper choice of tail size and location of tail 
and of center of gravity of the missile. 

The second method which has been used to reduce 
bore-sight errors from angle of attack changes is to 
mount the intelligence device on trunnions and cou¬ 
ple it to the controls in such a manner that the ref¬ 
erence axis of the intelligence device is rotated to 
compensate for changes in angle of attack. This com¬ 
pensation can be made perfectly for steady-flight 
conditions, but there are residual errors under dy¬ 
namic conditions. In addition a stability problem 
arises, as in the case of computers, and usually only 
a partial compensation can be made if hunting trou¬ 
bles are to be avoided. In some instances, such a 
coupling of intelligence device to controls has been 
used as a means of damping, as previously discussed. 
In such a case the bore-sight error is not completely 
eliminated. 

A final source of bore-sight error is the inevitable 
inaccuracy in construction of the internal shape of 
the vehicle. The axis of zero moment, i.e., the direc- 




254 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


tion of travel, will vary somewhat from one missile 
to another of the same intended shape. Such errors 
can be measured in a suitable wind tunnel, but only 
with great expenditure of time, and the requirements 
for uniformity of flow in the wind tunnel are difficult 
to meet. In the case of Pelican and Bat missiles, er¬ 
rors of about 0.8 degree were found from this source. 

With sufficient care and exclusive of actual equip¬ 
ment failures, it should be possible to reduce the bore- 
sight error to the order of magnitude of 1 degree. 


12* SENSITIVITY, LAG, AND 

MINIMUM RANGE 

The preceding discussion of pursuit curves and 
bore-sight errors has assumed certain ideal charac¬ 
teristics of the intelligence device and has not taken 
into consideration (1) that the intelligence device 
must look away from the target to produce error 
signals and (2) that it has a certain lag. 

Let us consider first an intelligence device without 
lag which gives an output signal proportional to the 
error angle up to a certain angle a m and which gives 
a constant output for angles greater than a m . The 
maximum output signal gives the minimum radius of 
curvature of the path p m . For any error angle g 
smaller than a m it will be assumed that the radius of 
curvature p c is given by p c /p w = <r m /<7; i.e., the radius 
of curvature is infinite for g = 0. If then the discus¬ 
sion of the pursuit curve is reexamined, it is seen that 
the axis of the intelligence device must point away 
from the target by an angle g when the radius of 
curvature is p c . 

An exact solution of this problem is difficult and 
has not been carried out. We have already seen that 
the pursuit-curve error is entirely analogous to a 
bore-sight error of magnitude tan -1 (^ sin 0 O )/(1 + n 
cos <f> o), where n is the ratio of the apparent target 
speed to the speed of the missile and </> 0 is the azimuth 
of the missile referred to the direction of apparent 
motion of the target. This equivalent bore-sight error 
varies during the flight of the missile along a pursuit 
curve because <t> 0 varies, and the pursuit-curve calcu¬ 
lations are complicated by the determination of <f>o as 
a function of the initial azimuth, initial range, and n. 
It will be instructive, however, to consider the sim¬ 
pler problems of the effect of the value of a m on the 
miss arising from a constant bore-sight error e. 

This problem has been worked out by Skramstad 
(unpublished) for the case where e is small, the lateral 
displacement y of the missile from a line joining the 


release point to the final target position is small com¬ 
pared with the initial range x 0 — x, and the slope of 
the trajectory dy/dx to the same line is small com¬ 
pared with 1. The differential equation is found to be 

Pc dX Pm&m Pm(TmL\X Xq / (XX _| 

The solution of this differential equation was found 
to be 

/ \ 1 X 0 H ( Xq /Pm^m) ^1 | 

y = e(x 0 - x) log--- + e 

L Xq X Xq/ p m G 7n J 

dy ^ ^ j J~ Xq _ H (Xq/ p m (J m ) ~1 

dx Lxq X Xq/ PmGm _J 

d 2 y = -e 

dx 2 Xq — X 

where 


H{x) = 1 + l + i + ^ + ••• 

For practical values of x 0 , p m , and G m , the second term 
in the brackets is entirely negligible. If, at a range 
x 0 — x = R m , the missile is assumed to follow a circle 
of radius p m , the miss M is given by 


Whence 


M -y + *4i 


1 

2 p m 


M = t{R m + pmOrn) 

" Pm 

The range R m at which g = G m or p c = p m is equal 
to p w €, 

since 

J_ = d/y = _ 6 = -j 

Pc dx 2 Xq — X R 

Hence the miss is given by 


M = p m G m e + |p m e 2 


= 1/15 we obtain the following 


If we assume G m 
values: 


0.05 
0.10 
0.15 
0.20 
0.25 

The effect of the value of 
if the missile travels 


M 

Pm 

0.0046 

0.0117 

0.0213 

0.0333 

0.0480 


r TO is zero if e = 0. Thus, 


a collision course and the bore- 











STRENGTH PROBLEMS OF HOMING MISSILES 


255 


sight error is zero, there is no error from limited sen¬ 
sitivity of the intelligence device. When e is not zero, 
the effect of a m is to give an error term proportional 
to €, whereas the miss due to bore-sight error alone 
is proportional to c 2 . Hence the effect of a m predomi¬ 
nates at small values of e. 

It is possible to use the formulas previously given 
to compute the pursuit-curve errors by means of an 
equivalent bore-sight error for the azimuth angle </> 0 
of the missile, referred to the direction of the appar¬ 
ent position of the target, changes from <£ 0 to 
<t> = 0o — dy/dx. By introducing the value of dy/dx 
in the expression for equivalent bore-sight error, 
there is obtained 

sin ( 0o — e + e log — + e log - ) 

€ __ _V_ Pm _ € / 

n + cos (<f>o — e + e log — + e log 

\ Pm € / 


This can be solved for RJ and introduced in the 
equation previously given to find the corrected miss. 
For most cases the total equivalent bore-sight error 
may be approximated in practice by 



The magnitude of the additional error is not large. 

If the intelligence device overloads as the target is 
approached, the directional information may dis¬ 
appear completely, or the error signals may even be 
reversed in sign. Furthermore, certain types, such as 
radar, have a minimum range within which no error 
signals are given. If the minimum range is R m , the 
missile in this case continues in a straight line and the 
miss is given by 

M = y + = e{R m + p m <T m ) 


This transcendental equation can be solved for e as 
a function of fa, n and x 0 /p m , and with the expression 
for M in terms of p m , a m , and e the values of M/p m 
as a function of a my <t> 0 , n, and x Q /p m can be found. 
The accuracy is not very good when \(j> — <f> 0 \ ex¬ 
ceeds about 0.3. 

Since large pursuit-curve errors correspond to 
large values of €, the term in a m is the smaller, but 
for u m = 1/15 corresponds to increasing the miss by 
50 to 100 per cent. 

The effect of a time lag r in the intelligence device 
may be estimated in a qualitative way by consider¬ 
ing this effect also as an equivalent bore-sight error. 
The error signal actually present at a range R is that 
appropriate to a range R + Vt, where V is the speed 
of the missile. This may be regarded as caused by 
an error angle at range R differing from that actually 
present by an equivalent additional bore-sight error 
equal to the change in the error angle between 
R + Vt and R. This is readily seen to be 

( 1 1 ) PmVmeVr 

R R + Vt) - R(R + Ft) 


The total equivalent bore-sight error is then 


[ 


1 + 


Pm<JmV ' 


R{R + Vt) 


] 


The range at which p c = p m is then given by 

7? ' - Jl 4- pm(JmV r 1 

Km + R m '(RJ + Fr)J 


For example, if R m = 2,000 ft, as was found in 
one design of intelligence device, p m = 10,000 ft, 
<r m = 1/15, and e = 1/30, the miss is 89 ft. If R m is 
reduced to 200 ft, the miss is 29 ft, of which 7 ft is 
associated with the minimum range and 22 ft with 
the radar sensitivity a m . 

129 STRENGTH PROBLEMS 

OF HOMING MISSILES 

The structural design of a missile is in most re¬ 
spects entirely analogous to that of an aircraft. It is 
desired here to call attention to only two design con¬ 
ditions peculiar to missiles. 

For rocket or jet-propelled missiles it is often in¬ 
efficient to design the power plant of sufficient size 
to include the take-off condition. It is better to use 
assisted take-off by means of a catapult or by special 
launching rockets. In German experience with anti¬ 
aircraft and long-range missiles, accelerations from 
2 g to 16*7 have been used in launching. The missile 
structure must be designed to withstand the inertia 
loads produced, and so must all the parts of the intel¬ 
ligence device and control system. 

The second design condition peculiar to missiles is 
that of full control application at maximum speed. 
Because of the bore-sight error which may be pres¬ 
ent, the intelligence device will call for the maximum 
rate of correction. The missile must be able to with¬ 
stand the loads so produced, or some automatic de¬ 
vice must be used to limit the acceleration Avhich may 
be imposed. 











256 


SOME ASPECTS OF THE DESIGN OF HOMING AERO-MISSILES 


12.10 LAUNCHING PROBLEMS 

OF HOMING MISSILES 

Some discussion has been given of launching prob¬ 
lems in the discussion of the relation between the field 
of view and the permissible motions of the missile. 
The target must come within the field of view and 
remain there. A sighting device or computer may be 
required for determination of the proper release time. 

A special problem encountered in the release of 
missiles from aircraft is that of avoiding maneuvers 
which cause contact between missile and aircraft 
after release. It appears essential to keep the homing 
device disconnected from the controls until the mis¬ 
sile is at a safe distance from the aircraft. The choice 
of airspeed and attitude at release can be made after 
the study of computed trajectories confirmed by ex¬ 
perimental tests. In general the airspeed must be 
lower than the horizontal flight speed of the missile, 


the amount by which it should be lower depending 
somewhat on the drag-weight ratio of the missile. In 
some cases interference effects between the missile 
and the aircraft may cause an unfavorable trajectory. 
However, no such effects have been encountered for 
missiles of high wing loading. 

12.11 CONCLUSION 

A survey has been given of some aspects of the de¬ 
sign of homing missiles for flight through air, pri¬ 
marily to place on record that part of the experience 
in the study of the radar-homing glide bombs, Peli¬ 
can and Bat, which is likely to be of value in the 
future development of homing missiles. None of the 
topics have been treated in detail in this chapter, but 
it is believed that the discussion is sufficient to indi¬ 
cate the nature of the problems and possible methods 
for their solution. 





Chapter 13 

THE DESIGN OF HIGH-ANGLE DIRIGIBLE BOMBS 


131 DESIGN PROBLEMS 

1311 General Considerations 

I N general, the structural configuration and di¬ 
mensions of high-angle dirigible bombs are suffi¬ 
ciently restricted by tactical considerations to pre¬ 
clude any extensive liberty in their design. Except 
for exceptional cases, they must be carried within 
the bomb-bays of airplanes which are designed pri¬ 
marily for carrying standard bombs comprising only 
the explosive case and sufficient tail-fin area to pro¬ 
vide reasonable weathercock stability of the bombs 
in flight. Thus the bomb-bays, including the carrying 
racks, are designed to pack a maximum number of 
compact explosive cases, with very little additional 
space for anything except the most rudimentary type 
of fin surface at the tail. Moreover, there existed an 
almost unsurmountable resistance on the part of the 
military to accept a dirigible bomb design which 
would not pack in existing bomb-bays in numbers 
equal to the number of standard bombs normally car¬ 
ried. This point of view will undoubtedly be changed 
to a reasonable degree when in the future such special 
weapons have proved their comparative worth in 
terms of potential effectiveness. However, during the 
recent period the design of special bombs was defi¬ 
nitely handicapped by this understandable but none¬ 
theless disconcerting viewpoint, which necessarily 
dominated developments. 

A few simple basic principles are involved in the 
aerodjmamic design of dirigible bombs, and the dis¬ 
cussion which follows is intended to outline in the 
simplest possible nontechnical terms certain major 
considerations which suffice to cover the broad as¬ 
pects of the problem. 

Control of a bomb in flight requires first of all con¬ 
trol of forces at right angles to the direction of flight 
so that the bomb may be deviated sideways, in any 
direction, from its normal trajectory. Since the dir¬ 
igible bombs contemplated here do not include the 
use of any internal power source such as jets, there 
remain only aerodynamic forces to accomplish the 
maneuvering of the bomb. Aerodynamic forces imply 
lift surfaces of adequate area to generate the desired 
forces, and means whereby these lift forces may be 
controlled at will. As to the method of obtaining lift 


forces from such surfaces as may be placed on a high- 
angle bomb, dimensional limitations leave no alter¬ 
native but to yaw or pitch the entire bomb so that it 
attains a trim angle of attack in the airstream. Since 
the lift force increases in proportion to the angle of 
attack, it is desirable to make it as large as possible, 
but drag forces increase disproportionately and soon 
reach the point where the ballistics of the bomb be¬ 
come intractable, so a compromise is required. More¬ 
over, the rudder surfaces start stalling seriously at 
around a 20-degree angle of attack, so that other 
problems are introduced if very large y aw angles are 
contemplated. Experience indicates that it is not de¬ 
sirable to maintain trim angles in excess of 15 degrees, 
and hence all dirigible bombs designed in this project 
have been held closely to this value. 

131,2 Mass vs Maneuverability Relations 

The lift force generated by any aerodynamic sur¬ 
face having an angle of attack in an airstream at 
velocities below the critical velocity at which the air¬ 
flow regime changes because of compressibility ef¬ 
fects may be expressed as: 

V 2 

Lift force per unit area = Clp~^~ 

0r . y 

Lift force = Area • Cl • p~n 
z 

where V = velocity of the airstream: 
p = density of the air; 

Cl = a numerical constant characteristic of 
the aerodynamic surface. 

The pV 2 /2 term is, of course, something over which 
we have no control, but it is of interest in the case of 
high-angle dirigible bombs because, in spite of the 
ever-increasing velocities of a falling bomb, the 
V 2 factor results in a practically constant maneu¬ 
verability. That is, if control is applied continuously 
throughout the drop, the radius of curvature of the 
resultant deviation path in space is substantially a 
constants This arises from the fact that, whereas the 
total velocity of the bomb along its trajectory is 

a The radius of curvature of the VB-1 is about 27,000 ft; 
for the VB-2, approximately 32,000 ft; and for the VB-3, about 
19,500 ft. 


257 



258 


THE DESIGN OF HIGH-ANGLE DIRIGIBLE BOMBS 


steadily increasing under the constant acceleration of 
gravity, the lift or deviation force is increasing as 
V-. Geometrical considerations will show a resultant 
deviation of practically constant radius. Thus, to the 
extent that the fin surfaces retain their effectiveness 
at high velocities, the maneuverability of a dirigible 
bomb will be practically constant anywhere in its 
flight. This, of course, is a very pertinent character¬ 
istic. 

The lift coefficient Cl is a dimensionless constant 
describing the effectiveness of the bomb body and its 
fins in producing aerodynamic lift. Its numerical 
value is determined by actual measurements, either 
by wind-tunnel tests on scale models or by actual 
drop tests of full-size bombs. However, here we are 
concerned with the broad fact that lift forces gen¬ 
erated aerodynamically are proportional to the area 
of the airfoil. This is of interest in verifying an in¬ 
tuitive guess that the more fin area applied to a bomb, 
the greater the lift or maneuverability, but it is more 
important in indicating that the size of bomb which 
may be maneuvered effectively with a given type of fin 
configuration is definitely limited. 

In considering the maneuverability of a bomb we 
are concerned with the rate at which it is deviated 
from its normal path when control is applied; that is, 
maneuverability depends upon the sideways or de¬ 
viating acceleration. Since a = F/m, the deviation 
acceleration is equal to the total lift force divided by 
the mass of the bomb. From the above, since the lift 
force is proportional to the area L 2 , and since mass is 
proportional to L 3 , it is found that, for bombs of 
similar design and density factor, deviation accelera¬ 
tion is proportional to L 2 /L 3 , or 1/L. 

Thus for a particular design of bomb and fins, the 
maneuverability of the bomb decreases in proportion 
to the increase in lineal dimensions, if the density 
factor remains constant and the dimensions of all 
parts simply increase proportionately. In terms of 
weight, the decrease in maneuverability with in¬ 
creased size is shown in the following table. 



Lineal 

Maneu¬ 

Inertial 


Wt. of 

dimension 

verability 

moment 

Period of 

bomb 

factor 

factor 

about c.g. 

oscillation 

1,000 

1.0 

1.0 

1.0 

1.0 

2,000 

1.26 

0.79 

3.18 

1.26 

8,000 

2.0 

0.50 

32.0 

2.0 

12,000 

2.29 

0.44 

63.0 

2.29 

22,000 

2.8 

0.36 

172.0 

2.8 


Hence, if it is assumed that a 1,000-lb bomb case 
has been supplied with a controllable fin structure 


whose area and aerodynamic effectiveness are such 
as to produce adequate lift for a certain desired ma¬ 
neuverability, then the same design in a 2,000-lb 
bomb will have only 79 per cent as much maneuver¬ 
ability as the smaller unit. In the larger sizes such as 
the 12,000-lb, the loss in maneuverability to 44 per 
cent of the 1,000-lb unit becomes very serious and for 
really satisfactory performance would require an in¬ 
crease in lift-surface area amounting to about 2.25 
times the area provided by merely scaling the 1,000- 
lb bomb to the larger size. This is a fundamental 
difficulty in designing dirigible bombs which pro¬ 
hibits unlimited expansion into large sizes without 
major modification of the design. In particular, it re¬ 
quires an ever-expanding area or lift surface to main¬ 
tain a fixed maneuverability. This circumstance must 
be borne in mind carefully in considering the applica¬ 
tion of the dirigible bomb principles to larger and 
larger bombs. The current development of a 12,000- 
lb bomb (Tarzon) is illustrative, for after using all 
available space for the most efficient lift surface dis¬ 
tribution practicable, the maneuverability is down 
to about 40 per cent of that obtainable with the 
1,000-lb VB-3. While this may still be quite satisfac¬ 
tory for azimuth steering, its adequacy for range 
steering is seriously questioned and will require actual 
trial to provide any satisfactory conclusions about 
the practicability of this bomb. 

13,1,3 Mass vs Yaw Oscillation 

Characteristics 

Another aerodynamic feature characteristic of high- 
angle dirigible bombs having broad implications in 
the application of controls to very large bombs per¬ 
tains to the yaw and pitch oscillations induced when 
control is applied. 

As already mentioned, the necessary lift forces are 
obtained by applying rudder or elevator to the con¬ 
trol surfaces, which yaw or pitch the bomb into a 
trim angle of attack of about 15 degrees, whereby the 
desired aerodynamic lift forces are brought into play. 
The application of the yawing moment by the rudder 
induces an overshoot of the bomb beyond its normal 
trim angle, and the steady-state trim angle is ap¬ 
proached only by a damped oscillation about the 
trim position. Unfortunately, the aerodynamic damp¬ 
ing induced by such oscillation is rather small; hence 
care must be taken in the application of rudder or 
elevator so as to limit the initial overshoot of the yaw 
angle as much as possible. It can be shown that, if 










DESIGN PROBLEMS 


259 


the small damping term is neglected, when rudder is 
applied instantaneously from neutral position, the 
initial yaw angle will be double the normal trim angle, 
and the yaw angle will be greater by far when the 
change involves immediate deflection from full rud¬ 
der in one direction to the opposite angle. It will also 
be evident that a very slow application of control will 
inhibit overshooting; in fact, no hunting will occur if 
the rudder application is in phase with the yawing 
angle. Thus the rate at which rudder is applied must 
be limited to a value somewhat comparable to the 
natural period of the bomb oscillation about its cen¬ 
ter of gravity. In the 1,000-lb VB-3 having periods 
ranging, because of ever-increasing velocity, from 
about 2.5 seconds at release to 0.7 second near the 
end of a 15,000-ft drop, experience has indicated that 
reasonable suppression of oscillation is obtained with 
rudder speeds of about 1.2 seconds for travel from 
neutral to full deflection. By way of comparison, the 
12,000-lb Tarzon design indicates a period of about 
1.75 greater than the VB-3; hence, the maximum rud¬ 
der speed must be reduced to about 2.0 seconds from 
neutral to full deflection. Moreover, this reasonably 
short period for the 12,000-lb bomb oscillation is 
achieved only by increasing the weathercock stabil¬ 
ity of the bomb relative to that of the VB-3, a com¬ 
promise which is obtainable only at a sacrifice in ef¬ 
fective lift for a given area of fin surface. The details 
regarding this inherent loss of lift with increasing 
weathercock stability are beyond the scope of this 
discussion. However, it is mentioned merely to indi¬ 
cate that the increased moment of inertia of larger 
bombs leading to embarrassingly long oscillation 
periods cannot be compensated for by indiscrimi¬ 
nately increasing the stiffness or weathercock stabil¬ 
ity by greater tail stabilization. It might appear that 
slowing up the rudder action arbitrarily to handle the 
long-period problem is sufficient. Such a device will, 
in fact, serve to keep the yaw oscillations within rea¬ 
sonable bounds. b However, it must be remembered 


b An alternative method might be devised whereby artificial 
damping is introduced by properly controlling the speed and 
phase of rudder applications, or by applying some kind of 
aerodynamic brake. While considerable thought has been given 
to this problem, it has not yet appeared amenable to solution 
without an excessive amount of auxiliary equipment. In addi¬ 
tion to added power requirements for special controls, it 
appears to require at least two or perhaps four additional gyro 
controls. Such devices might be found practicable for the very 
large bombs, such as the VB-13, but it is questionable whether 
the added complexity and accompanying decrease in reliability 
of the instrumentation would be warranted, except as a last 
resort. 


that the high-angle steering problem also demands 
reasonable speeds in response to controls; otherwise, 
accurate steering would be impossible. The optimum 
design is one of compromises throughout, and great 
care must be taken that the whole purpose of the 
control feature is not vitiated. 

The above discussion, while far from complete, is 
intended to point out forcefully that there are limi- , 
tations to the practicable weight of dirigible bombs, 
unless the space made available for the aerodynamic 
surfaces is allowed to expand disproportionately. One 
cannot cope with the area vs mass relationship, or 
the moment of inertia vs mass stability problem 
without some additional freedom in the design of the 
aerodynamic surfaces other than mere uniform ex¬ 
pansion of dimensions. 

1314 The Roll Torque Problem 

Dirigible bombs of the type discussed here all re¬ 
quire stabilization of the bomb in roll about its 
longitudinal axis. This basic requirement that the 
roll orientation be fixed throughout the flight of the 
bomb will be evident on considering the necessity of 
defining the plane of action of the rudders and ele¬ 
vators with respect to azimuthal and range coordi¬ 
nates in space. Thus the rudders must be maintained 
parallel to the vertical or azimuthal reference plane, 
and the elevators to a horizontal or range reference 
plane. The use of a suitable gyro control acting 
through roll-stabilizing ailerons has accomplished the 
required stabilization in the Razon-type bomb, VB-3. 

Throughout this chapter the importance of this 
roll control problem has been stressed, and particular 
attention has been called to the difficulties encoun¬ 
tered in attempting to,stabilize the flat-fin cruciform- 
type empennage. Indeed, reliable performance was 
never attained with this type of structure when it 
was subjected to simultaneous yaw and pitch con¬ 
trol, although attempts to “brute-force” stabiliza¬ 
tion with larger and larger ailerons were carried to 
the point at which practical considerations of power 
and space limitations called a halt to further at¬ 
tempts. At the same time, considerable effort was 
made to analyze the problem from a quantitative 
aerodynamic viewpoint by means of wind-tunnel 
tests designed to measure the roll torques generated 
when a bomb equipped with a cruciform-fin structure 
was exposed to simultaneous pitch and yaw. Such 
tests were made in several wind tunnels, but reliable 
quantitative information was found very difficult to 






260 


THE DESIGN OF HIGH-ANGLE DIRIGIBLE BOMBS 


obtain. However, numerical data were secured which, 
qualitatively, showed the typical roll torque charac¬ 
teristics that were anticipated from the early cursory 
visual consideration of the structure. Quantitatively, 
the data were of no great value. On the contrary, 
they were misleading to the extent that in some quar¬ 
ters they encouraged the belief that the roll torque 
problem was less serious than seemed evident from 
actual drop tests. In large measure, these deceptive 
data were responsible for the illusive hope that suffi¬ 
ciently large ailerons could be applied to the struc¬ 
ture to assure roll stabilization. 

Although this is not the place to discuss the aero¬ 
dynamic aspects in detail, the importance of the 
problem, in view of future work on dirigible missiles, 
warrants brief consideration of the source of the roll 
torque phenomenon. Indeed it seems imperative, for 
even now development work appears to be under way 
using cruciform-type empennage designs with too 
little consideration for the seriousness of the problem. 
It is as if this contractor had never been plagued with 
the roll problem or, if so, had labored under delusions 
regarding its seriousness. 

As regards the source of the roll torques developed 
when a flat-fin cruciform-type tail is subjected to the 
cross wind associated with combined pitch and yaw 
angles of attack, a cursory study of a model will pro¬ 
vide all the qualitative information required. 

The first and simplest consideration is that the 
structure comprises flat radial surfaces; hence, any 
unbalanced lifts produced on these surfaces represent 
roll torques whose magnitude is merely the product 
of the unbalanced force and the radial distance to its 
effective center of pressure. It is pertinent, then, to 
examine the conditions which will give rise to un¬ 
balanced lift forces. , 

If we assume that there are no interference effects, 
that is, no shading of one surface by another due to 
an asymmetrical wind, unbalanced lift forces (and 
therefore torques) will be produced, except when the 
direction of the wind vector is such that its projection 
in the plane of the lift surface is parallel to the axis of 
the empennage. In other words, any flat surface con¬ 
strained on a central axis vail always orient in a fluid 
stream so as to expose a maximum surface at right 
angles to the stream. It will be clear that, if a missile 
carrying a tail surface comprising a single flat fin is 
yawed in the windstream, it will roll on its axis until 
the plane containing the wind vector and the axis is 
normal to the surface; that is, it will roll until a 
maximum area is exposed to the wind. If such a 


missile is yawed in azimuth, it will be stable if the fin 
is vertical; but if given a pitch angle of attack, it will 
be stable in roll only if the fin is horizontal. In the 
case of an empennage having two flat fins at right 
angles, as in the cruciform structure, it will be evi¬ 
dent from similar considerations that it will be stable 
in roll if subjected to pure yaw or to pure pitch, but a 
combination of the two will develop a roll torque. 
This stability characteristic explains why the cruci¬ 
form structure is reasonably practical in the VB-1, 
which is controlled in azimuth only (or in range only, 
if desired), but is impractical for a Razon bomb 
steered in both range and azimuth. 

On the other hand, if the vertical and horizontal 
fins have unequal areas, the assembly will behave 
lik6 the single-fin structure to the extent that the 
two sets of fins are of unequal size. This effect was 
observed to a pronounced degree in actual bombs 
when an experimental attempt was made to use 
unequal fin areas to obtain greater maneuverability 
in range than in azimuth. Hence, the Azon empennage 
carries horizontal and vertical fins of equal size. 

The above discussion can be summarized by stat¬ 
ing that the roll torque characteristics of the cruci¬ 
form structure are inherent because the projected 
area of the assembly varies with the angle of roll; 
that is, the structure is not radially symmetrical as 
regards exposed area. Obviously, the area asym¬ 
metry could be improved by increasing the number 
of fins from, say, four to eight, but other practical 
considerations preclude such a design. Moreover, the 
simple explanation of roll torque considered thus far 
is only one part and not necessarily the major part of 
the roll torque problem. 

Thus far we have ignored the effect of interference 
or shading of one fin on another or of the body of the 
missile on the fin surfaces. We have also ignored the 
effects due to the shape of the body section of the 
empennage to which the cruciform fins are attached. 
The shape ol the body lying within the cruciform 
structure can markedly affect the roll stability pat¬ 
tern in both magnitude and angular distribution of 
the stability angles. Details regarding the latter 
effects can be obtained only from wind-tunnel studies. 

However, the shading effects may be considered 
profitably in a qualitative manner by recalling that 
any unbalanced lift forces on the radial fins of the 
cruciform structure will generate a roll torque. More¬ 
over, it is evident that any aerodynamic shading 
affecting the several fin surfaces unequally will neces¬ 
sarily result in unbalanced lift forces. 



DESIGN PROBLEMS 


261 


Consider a bomb equipped with a cruciform tail 
having elevator and rudder flaps on the trailing edge 
of each of the horizontal and vertical fins. In normal 
flight the tail section is subject to the wash of the 
bomb body, so that the inner portions of the surfaces 
may be exposed to very disturbed air while the fin 
areas at some greater radial distance outward may be 
in relatively smooth air. However, since all four fins 
are symmetrically disposed with respect to the cone 
of rough air behind the bomb body, and in normal 
flight the surfaces have no angle of attack, no roll 
torque of appreciable magnitude should develop. 

Now suppose that elevators are applied so that the 
bomb attains a trim angle of attack (15 degrees up¬ 
ward pitch for typical Razon bombs). The frontal 
area of the bomb body has now increased because of 
the attack angle, and the area of the disturbed wake 
will likewise increase, accompanied by even greater 
turbulence. Obviously, this will result in an increased 
shading of the upper half of the vertical fin, accom¬ 
panied by a decreased shading of the lower half; but 
since the vertical fin is still parallel to the wind 
stream (no rudders applied) no lift will be developed 
and no roll torque. Similarly, the right and left sec¬ 
tions of the horizontal fin are symmetrically disposed 
in the wake, and no torque should result. Thus, in pure 
pitch or yaw alone the bomb should be stable in roll. 

When the elevator angle of attack is maintained 
and right rudder suddenly applied, consider the aero¬ 
dynamic reaction due to the rudder application 
before the bomb has yawed to any appreciable extent 
in response to the rudders. It will be apparent that, 
as described in the preceding paragraph, the upper 
section of the vertical fin in the tail of the bomb (the 
nose of which is tilted upward 15 degrees in its ele¬ 
vator trim position) is heavily shaded within the 
wake of the bomb body, while the lower section is in 
smooth air. Since the rudder flaps on the vertical fins 
are deflected, the lift force generated will appear as a 
thrust directed to the left on the vertical fins. How¬ 
ever, since the lower fin is more efficient aerodynam- 
ically than the shaded upper fin, the lift forces are not 
balanced, and the excess thrust on the lower vertical 
fin will generate a clockwise roll torque. As the bomb 
yaws into its proper trim position, it will be observed 
that a similar shading effect will be exhibited by the 
horizontal or elevator fins; but the resulting roll 
torque will be counterclockwise, tending to cancel the 
torque from the rudders. Thus, it would appear that 
if the yaw and pitch response of the bomb was always 
in trim with the rudder and elevator applications, 


roll torques due to shading effects might be relatively 
unimportant. 

However, yaw and pitch oscillations of a bomb are 
very poorly damped (the coefficient of damping is 
approximately 0.25 for the Razon bomb) so that, 
because of randomly imposed control applications, it 
is easy to induce transient angles of attack 1.5 or even 
2.0 times the normal trim angles. Moreover, as a 
result of such oscillations, it is possible that the angle 
of attack of the bomb may at times be the opposite 
of that called for by the rudders or elevators. Should 
this occur, then the shading effects on the rudder and 
elevator will not cancel as in the case considered 
above; their roll torques will be additive, and the 
resultant magnitude will depend only upon the asym¬ 
metry in the lift forces induced by shading. It should 
also be remembered that this roll torque is a phe¬ 
nomenon due to control-flap action coupled with 
aerodynamic shading effects and is distinct from and, 
depending upon the phase of the events, ipay be 
additive to the torques due to radial asymmetry of 
the fin structure discussed earlier. It is also likely 
that it is the more important of the two. 

The qualitative description of these sources of roll 
torque, characteristic of the radial-fin cruciform 
structure, indicate conclusively that the torques are 
strongly impulsive in character. But, however short 
their duration, the energy will be exhibited as a roll 
acceleration of the bomb, which must be balanced by 
an equal and opposite acceleration produced by the 
roll-stabilizing ailerons, which for practical reasons 
are necessarily limited in capacity. It will also be ap¬ 
parent that the magnitude of the transient roll 
torques will depend to a large extent upon the for¬ 
tuitous coincidence of certain combinations of yaw or 
pitch oscillations together with control applications, 
with the result that the occurrence of excessive 
torques will be a random event. Both the impulsive 
and unpredictable occurrence of the phenomenon 
have been observed repeatedly in actual drop tests. 

Finally, it may be said that the magnitude of the 
torques due to control asymmetry resulting from 
aerodynamic shading on a cruciform structure will be 
directly proportional to the lift forces produced by 
the control, to the area of the disturbed wake in 
which the fins are emersed, and to the degree of tur¬ 
bulence within the wake which gives rise to the dele¬ 
terious aerodynamic effects resulting in asymmetrical 
lift forces. It wall be clear, therefore, that at low 
velocities the net effect may be small, but with in¬ 
creasing velocity the disturbed wake behind the 



262 


THE DESIGN OF HIGH-ANGLE DIRIGIBLE BOMBS 


bomb body will increase in diameter and in degree of 
turbulence, with a resultant rapid increase in the 
aerodynamic asymmetry of the fin surfaces. Thus, 
failure of roll stabilization may not occur until the 
velocity of the missile attains a critical value. Doubt¬ 
less the failure of wind-tunnel studies to produce a 
reliable quantitative analysis of the roll torque prob¬ 
lem consistent with the phenomena actually ob¬ 
served in drop tests is due to the lack of wind-tunnel 
tests at sufficiently high velocities. It is a well-known 
fact that it is quite impossible to extrapolate low- 
velocity observations involving especially the type of 
phenomena referred to here as shading since the flow 
regime undergoes marked changes beyond certain 
critical velocities which are a function of the aero¬ 
dynamic shape of the structure. 

Indeed this effect has been clearly observed in 
nose-camera motion pictures wherein, occasionally, 
bombs of the VE^-l type carrying cruciform fins have 
been observed to develop spurious trim angles of 
attack up to 5 degrees or more near the end of a 
15,000-ft drop, even though no control was applied. 
This is a direct result of the blanking of the stabiliz¬ 
ing fins by the turbulent wake of the bomb body. 
Similar effects have been observed visually in the 
last several thousand feet of 15,000-ft drops of stand¬ 
ard 1,000-lb bombs. In these cases the effect was 
observable because of a slow rotation deliberately 
imposed by very small ailerons on the standard 
bombs for the purpose of eliminating dispersion. The 
effect was exhibited occasionally by the sudden initi¬ 
ation of a conical wobble of the bomb and an accom¬ 
panying visible increase in trail angle compared with 
others in the group. 

Two obvious sources of roll torques in radial fin 
structures have been considered. There are more 
complex effects, such as mutual interference of the fin 
surfaces themselves, that should be considered. How¬ 
ever, this very qualitative and cursory discussion 
should serve to emphasize that the use of radial-fin 
structures on dirigible missiles should be contem¬ 
plated with caution. True, the shading effects may be 
ameliorated by the use of far better streamlining of 
the aft portion of the missile than was possible in the 
design of the high-angle dirigible bombs. Moreover, 
the roll torques due to all causes fade out rapidly as 
the angles of attack required for adequate maneuver¬ 
ability are reduced below the large values necessary 
in the dirigible bombs. But no empennage remains 
well-streamlined when large angles of attack are in¬ 
volved; and if these angles are kept small, maneuver¬ 


ability must be sacrificed in proportion unless larger 
lift surfaces are permissible. But if it is assumed that 
cleaner design of the empennage may be attained 
than was possible in converting to a dirigible type the 
aerodynamically inept standard bomb, it must be 
remembered that such gains may, to a large extent, 
be vitiated at the very high velocities contemplated 
for some of the proposed highly maneuverable missiles. 
Hence, a careful study of the problem is warranted. 

The Division is well aware of the compactness and 
particularly the attractively “clean” aerodynamic 
features of the radial-fin structures which recommend 
them, particularly for high-speed missiles. In the case 
of the dirigible-bomb project, time did not permit an 
exhaustive study of the roll torque problem and this 
definitely precluded the use of radial fins in the 
Razon-type bomb. The alternative was immediate 
adoption of a type of structure inherently free of the 
problem. Thus, the use of cylindrical shroud surfaces 
eliminates any possibility of a serious roll torque prob¬ 
lem, since such surfaces have radial symmetry and 
are so disposed that any lift forces produced act ra¬ 
dially and cannot generate a roll torque. Hence, from 
the roll torque aspect the structure is inherently free 
even of the aerodynamic shading problem. The prac¬ 
tical requirement of suitable radial supports for the 
shroud apparatus does introduce some radial asym¬ 
metry. Likewise, simplicity of design calls for a modi¬ 
fication of the ideal circular shroud to an octagonal 
shape which facilitates the use of simple control flaps. 
But these modifications result in only minor asym¬ 
metry and present no serious problem in roll stabili¬ 
zation. For use on missiles whose velocity range is 
within the permissible limits for this type of struc¬ 
ture, their adoption may be contemplated without 
serious worry about the roll stabilization problem, 
and reliable performance may be assumed with confi¬ 
dence. This obvious advantage of the shroud-type 
control and lift surface over the radial-fin type for 
highly maneuverable missiles warrants some exten¬ 
sive study to determine the practical limit of velocity 
for which they are applicable. 

13 2 FUTURE APPLICATIONS 

OF DIRIGIBLE BOMB TECHNIQUE 

Alleviating Limitations 
on Size of Dirigible Bombs 

The above discussion is intended to outline the 
salient features of high-angle dirigible bomb design 



FUTURE APPLICATIONS OF DIRIGIBLE BOMB TECHNIQUE 


263 


which impose inherent limitations on the weight of 
bombs which may be controlled in the manner of the 
VB-1, -2, and -3 units. It is indicated that in the 
12,000-lb VB-13 we have already passed the point at 
which truly effective control is obtainable—effective 
in the sense that these bombs are intended to permit 
accurately directed maneuvers of 1,500 to 2,000 ft in 
order to secure hits on targets executing evasive 
action. 

These considerations lead to the inevitable conclu¬ 
sion that specifications for very large dirigible bombs 
must allow for a disproportionate increase in the size 
of the aerodynamic surfaces as compared with the 
1,000- to 2,000-lb sizes. Alternatively, the use of some 
type of controlled jet-power units must be contem¬ 
plated, provided adequate side thrusts are obtainable 
with compact units of a type which can be turned on 
or off as occasion requires. 


13 2 2 High-Velocity Dirigible Projectiles 

The success of high-angle dirigible bombs might 
lead one naturally to the conclusion that similar 
aerodynamic methods of control are applicable to 
other types of projectiles. Aside from the weight con¬ 
siderations mentioned above, there are also limita¬ 
tions on the velocity at which such aerodynamic con¬ 
trols may be reliable. The extent to which the velocity 
of the VB-3 may be increased without loss of control 
is not yet established by experience. It is known 
from high-speed wind-tunnel tests that the drag 
coefficients on all structures start increasing rapidly 
above a speed of about 800 ft per sec. This corre¬ 
sponds approximately to the terminal velocity of a 
15,000-ft drop. The VB-3 has been tested at 20,000 ft 
with no indications of failure, and it is felt that at 
25,000 ft it should be quite satisfactory. However, 
while the Division has no data on drag coefficients 
above the critical velocity mentioned above and 
hence cannot calculate the probable terminal veloc¬ 
ity, there is little reason to expect a terminal velocity 
much above 1,000 ft per sec with the small bombs. 
Because of lack of adequate information at present, 
the performance of the aerodynamic controls, either 
aileron or directional, is an unanswered question 
when the terminal velocity due to aerodynamic drag 
has been reached. This is especially true of the com¬ 
plete lift shroud-control shroud assembly of the 
VB-3 or VB-13 type. However, it will be recalled 
that this type of design was used because it was 


amenable to roll control whereas the more orthodox 
cruciform-fin design was not. While the latter design 
is undeniably a “cleaner” and more compact aero¬ 
dynamic structure and amenable to higher speeds, its 
use must be envisaged only with great caution in 
regard to its unfavorable characteristics. 

In addition to the roll control problem introduced 
by the use of the cruciform fins, certain additional 
features should be mentioned. It will be recalled that 
one of the attractive features of the cruciform-fin de¬ 
sign is its compactness as ordinarily fitted into a 
bomb-bay. As applied to the standard 1,000- and 
2,000-lb bomb cases in the VB-1 and -2 designs, only 
about half of the fin area projects beyond the diam¬ 
eter of the bomb body. It is of interest that a success¬ 
ful cylindrical or octagonal control shroud for these 
bombs cannot be made unless its diameter is greater 
than the bomb diameter. The standard bomb cases 
used in these bombs have such a short tail that no 
real streamlining of the aft portion is possible, and 
even at low velocities normal airflow breaks away 
behind the main section of the bomb. At very high 
velocities one would expect a complete break-away at 
this point, with a resultant decrease in effectiveness 
of the airfoil within this diameter. Low-speed tunnel 
tests cannot show this effect, but the Division has 
direct evidence, from trajectory photographs, of the 
cruciform-type VB-1 which shows that near the end 
of flight these bombs evidence changes in their aero¬ 
dynamic behavior. Similarly, in the case of standard 
bombs with standard tail fins, the evidence shows 
that they have lost some of their normal weathercock 
stability. This experience clearly indicates that if a 
cruciform-type fin structure is used, it should be 
placed on a well-streamlined conical section having 
only a gradual taper from the bomb body, in order to 
minimize break-away of the airstream and particu¬ 
larly to minimize shadowing of the fin structure by 
the bomb body when the bomb is yawed. 

It is believed that both the roll torque and shadow¬ 
ing effects discussed above are sufficient to justify 
discarding the cruciform-fin structure from any con¬ 
sideration for projectiles having the general propor¬ 
tions of the VB-1, -2, or -3 types, particularly if the 
aerodynamic control requires yaw angles of the order 
of 10 to 15 degrees. It is felt that these structures will 
be unreliable for such projectiles at velocities in ex¬ 
cess of about 800 ft per sec. With a longer tail section, 
providing possibilities of streamlining the empen¬ 
nage better, these objections might be materially 
mitigated, especially if only small angles of yaw or 




264 


THE DESIGN OF HIGH-ANGLE DIRIGIBLE BOMBS 


pitch are involved, but the problem deserves very 
serious study in any case. 


13 2 3 Jet Control of Bombs or Projectiles 

The use of aerodynamic lift surfaces on very large 
bombs entails inherent limitations as to the weight of 
the bombs, unless the area of the aerodynamic sur¬ 
faces is increased proportionally. Questions regarding 
velocity limitations also require study. Moreover, as 
high-velocity bombs are intended to attain deep 
penetration on impact, the effect of a 15-degree or 
greater angle of attack at impact must be considered. 
In using the VB-3, for example, there were observed 
a few cases in which the bomb had control applied 
just before impact, and the bomb deflected upward 
after entering the ground. Thus, any use of such 
bombs contemplating a semi-armor-piercing type 
requires further study. 

These problems all lead to the consideration of 
some means other than aerodynamic forces to achieve 
control of the trajectory. Moreover, certain contem¬ 
plated types of high-speed controlled projectiles re¬ 
quire a degree of maneuverability hardly amenable 
to an aerodynamic solution. The only evident alter¬ 
native at present is some type of jet power, applied 
radially without introducing roll torques. While the 
contractor is not sufficiently informed regarding 
either rocket or fuel-injection rocket-type jets, it 
would seem that the latter have attractive possibili¬ 
ties for controlled missiles. 

Rocket or jet-propelled missiles have been built in 
which control has been applied by means of deflectors 
in the high-velocity jet stream, resulting in a change 
in direction of the thrust. If this thrust is applied aft 
of the aerodynamic center of pressure, then the com¬ 
ponent of thrust normal to the axis which is induced 
by the deflectors serves to yaw the missile into a trim 
angle of attack, with a resultant change in course due 
to aero-dynamic lift. But the maneuverability ob¬ 
tainable by aero-dynamic forces is distinctly limited 
by practical limitations in surface dimensions. If 
mechanical problems can be met, it seems not im¬ 
practicable to apply the lateral component of the jet 
thrust at or perhaps slightly forward of the center of 
gravity, in which case the full benefit of jet thrust can 
be used to deflect the course of the missile, aided by 
the incidental aerodynamic lift induced by any yaw¬ 
ing of the bomb since this lift will be in the same 
direction as the jet thrust. 


Thus it would seem that the ideal controlled mis¬ 
sile would be propelled by an annular jet slightly for¬ 
ward of the center of gravity of the missile. This jet 
would be directed rearward at an angle of 45 degrees 
or less from the axis of the missile, depending upon 
the degree of maneuverability desired. If this jet is of 
the fuel-injection type, so arranged that the annular 
distribution of the fuel may be controlled, the direc¬ 
tion of the net thrust and hence the direction of 
flight would be controlled at will. Such control could 
apply a large fraction of the total thrust as a deviat¬ 
ing force. Moreover, if such a scheme could be de¬ 
veloped, the control of fuel distribution could be ac¬ 
complished with relatively simple, fast-acting, and 
compact control devices, well adapted to automatic 
control equipment. While such a self-propelled, con¬ 
trolled missile may appear rather fantastic, and, 
indeed, may be relegated immediately to the realm of 
the impractical on the basis of already known prop¬ 
erties of jets, other mechanical and performance fea¬ 
tures are very attractive. Initial acceleration to high 
velocity may be obtained by means of a rocket 
which could be released from the tail when expended. 
Above all, such jets appear to be the only means of 
circumventing the maneuverability limitations of 
aerodynamic methods for missiles requiring a very 
high degree of maneuverability, such as either 
ground-to-air or air-to-air antiaircraft weapons. 

A missile of the above type would be designed with 
a high degree of weathercock stability to minimize 
yaw. This would minimize the roll control problem, 
which is one of the really serious problems in all 
dirigible missiles. Alternatively, if the annular-jet 
power distribution is amenable to simple and rapid 
control, the missile may be allowed to rotate at a 
reasonable and controlled rate, with only the control 
apparatus stabilized in roll. Certain types of homing 
devices involving rotary scanning might make effec¬ 
tive use of such a scheme. 

These rather futuristic suggestions are offered pri¬ 
marily to stress the fact that the contractor’s four 
years of experience in the development of high-angle 
dirigible bombs has rather impressively indicated 
some of the limitations of normal aerodynamic solu¬ 
tions to the controlled missiles problem. In view of 
some of the weapons which have been visualized, it 
is considered that a substitute for the completely 
aerodynamic methods is imperative. The great 
weight of some of the contemplated bombs, the high 
velocity, and particularly the very high maneuver¬ 
ability desired in some of these weapons do not ap- 



FUTURE APPLICATIONS OF DIRIGIBLE BOMB TECHNIQUE 


265 


pear amenable to the usual aerodynamic controls. 
Fuel-injection rocket-type impellers to provide the 
necessary deviation forces appear to be the only 
solution. If developed to the point of practicability as 
a power source, such units will offer attractive possi¬ 
bilities for simplification of control mechanisms, es¬ 
pecially of the automatic or target-seeking type. For 
these reasons an intensive study of such methods not 
only is justified but is an urgently needed program. 
Indeed, if present airplane speeds are even margin¬ 
ally close to vitiating aerodynamic controls for anti¬ 
aircraft missiles, such methods are then already 
obsolete for future use. Yet some method must be 
devised whereby control of the missile between the 
launching point and target is made possible, however 
fantastic it may appear at this moment. 

13 2 4 High-Angle Armor-Piercing Bombs 

The current development work on the VB-13 
(12,000-lb dirigible bomb) is indicative of the impor¬ 
tance of controlling very heavy bombs of the semi- 
armor-piercing type. The difficulties in providing 
maneuverability without decreasing the available 
penetration have already been mentioned. Such 
heavy bombs require large control surfaces with re¬ 
sultant aerodynamic drags that seriously limit the 
terminal velocities. Moreover, aerodynamic controls 


require yawing of the bomb, which will seriously 
affect penetration; yet one cannot handicap the ac¬ 
curacy of control by prohibiting steering at the end 
of flight where it is most needed for precision control. 

As a solution to some of these problems, it has 
already been suggested that rocket accelerators 
might be used at the end of flight. It has occurred to 
the contractor that perhaps a shaped-charge bomb 
with a follow-through missile might be more prac¬ 
tical. It is suggested that a 36-in. diameter shaped 
charge about 48 in. long would penetrate 16 ft of 
concrete and leave an opening through which 200 to 
300 lb of follow-through explosive would be pro¬ 
jected. Such a bomb would also serve against steel 
armor. There remains the question about the influ¬ 
ence of impact angle on the effectiveness of the shaped 
charge. The follow-through charge would not appear 
to present insurmountable difficulties in view of the 
solution of a similar smaller-scale problem. 

A bomb of the above type would be smaller than 
the VB-13 and hence more adaptable to the dirigible 
type. It would have a penetrating power on some 
targets equal to the VB-13 and could be more de¬ 
structive for certain purposes. Since the only evident 
problems are the two mentioned above, both of which 
should be readily solved without a prohibitive 
amount of research, it would seem that such a bomb 
is well worth investigation. 



Chapter 14 

PROJECT ROC IN RETROSPECT 


“i INTRODUCTION 

T he present chapter was prepared for the pur¬ 
pose of collecting some of the lessons and experi¬ 
ences that were learned by a group engaged in the 
development of a family of guided missiles designated 
Roc. 

Chapter 12 covers some of the broad aspects and 
fundamental philosophy of the design of guided fly¬ 
ing missiles with particular reference to experience 
gained with projects Bat and Pelican. Since most of 
the conclusions and recommendations presented 
there express the corresponding sentiments of the 
Roc group most admirably, there is no need to re¬ 
capitulate the entire subject matter here. The present 
note will therefore be restricted to such supple¬ 
mentary remarks as appear appropriate to convey or 
preserve incidental bits of experience which were 
peculiar to Roc, to its developmental history, or to 
the personnel engaged in the project. 

Of the five types or categories of missiles enumer¬ 
ated by Dryden, namely, (1) autopilot-launched, 
(2) suicide-pilot directed, (3) remote-vision guided, 
(4) beam-borne, (5) self-homing, Roc started out as a 
project of type 4 (beam-borne); it was then devel¬ 
oped toward type 5 (self-homing) and finally built 
as type 3 (remotely guided), adapted to a variety of 
intelligence devices. While the advantages of versa¬ 
tility under conditions of rapid changes of the for¬ 
tunes of war are impressive, it must not be over¬ 
looked that the price of versatility is compromise of 
performance. The best device for any particular pur¬ 
pose is likely to be the one specially designed for it 
and nothing else. The best procedure, however, is not 
necessarily to lay down a utility specification and 
then have someone else meet it, but rather to make a 
careful study of the prospective merits and penalties 
of variations of any tentative specification. 

Many viewpoints of a technical, tactical, and 
strategical nature have to be reconciled. This requires 
the close cooperation of specialists in many fields, 
who should have some understanding of the realms of 
neighboring fields. The necessity for intimate coor¬ 
dination at the research and design levels has already 
been stressed in Chapter 12. Equally close coordina¬ 
tion is needed between design and preliminary re¬ 


search on the one hand and flight testing on the 
other. In fact, flight testing might well be considered 
more as a part of the research activity than as a 
function of proof testing, involving as it does 
technical problems of considerable magnitude in 
itself. 

Flight testing in the Roc project was the responsi¬ 
bility of the same organization as research and de¬ 
sign. The advantages of this union were keenly and 
frequently felt. So was (while it lasted) the relative 
proximity of the testing range to the location of the 
engineering organization. 

Guided missiles have been classified according to a 
great variety of viewpoints. One of these is that of 
the trajectory slope. Distinction can be made be¬ 
tween (1) bombs falling essentially in a ballistic 
trajectory corrected only by a small fraction of the 
height, (2) bombs beginning to fall in a ballistic 
trajectory but possessing aerodynamic devices to de¬ 
flect them appreciably from a passive trajectory 
through a distance commensurable with the height, 
and sufficient to make an accurate bombsight ap¬ 
proach unnecessary, (3) glide bombs which develop 
enough lift to carry their own weight in a steady 
glide and travel a distance several times the height of 
release, and (4) propelled missiles capable of climbing 
to considerable altitudes. 

Of these, the Roc project was deliberately assigned 
to fill category (2), the moderately steep dive bomb. 
Tactically, it is closely related to category (1), the 
steep high-angle bomb, from which it differs essen¬ 
tially by the concession of a wing system which ex¬ 
ceeds the caliber-circumscribed box so that more lift 
can be generated than by any system wholly con¬ 
tained within the dimensions of such a box. 


142 WING SYSTEM 

When a wing system is to be attached to the bomb 
body, the problem arises of how it should be ar¬ 
ranged. One school of thought favors an arrangement 
after the manner of an airplane, with the wings in 
one plane generating lift normal to that plane only; 
this requires banking of the vehicle into such an 
attitude as to tilt the lift into whatever direction is 


266 


WING SYSTEM 


267 


desired. This is the natural configuration for flat¬ 
flying or gliding bombs, which have to develop lift to 
balance gravity at all times and may require only 
secondary and occasional corrections in other direc¬ 
tions. 

On the other hand, missiles designed to fly steep 
paths (or, for that matter, those required to generate 
lift amounting to multiples of the path-normal com¬ 
ponent of gravity) may be called upon to develop 
lift quickly in any direction. This may demand large 
and rapid changes of roll attitude of a monoplane 
vehicle and therefore require a powerful aileron sys¬ 
tem. In a conventional airplane, the span and hence 
the available aileron leverage are large, the roll 
damping is appreciable, and the moment of inertia 
against rolling is relatively small. If, however, the 
wing system of a missile is of small span compared to 
the radius of gyration, and if the aileron loading is of 
necessity high, as in a dive bomb, then it is more 
difficult to generate and control large rolling mo¬ 
ments to initiate a rolling motion and to stop it in 
time. This problem becomes particularly severe when 
the lift demand drops to or near zero, only to increase 
again in the opposite direction without much of a 
lateral component present. It is here that the advan¬ 
tage of providing wings in all directions makes itself 
felt, and this is why arrangements of three radial 
wings at 120 degrees to each other and of four cruci¬ 
form wings at 90 degrees were tried and a universally 
jointed ring wing finally adopted in project Roc. The 
extra drag of the wing component not needed at any 
one time can usually be tolerated on a gravity-pro¬ 
pelled device. As to the extra apparatus, it is a ques¬ 
tion of weighing it against the more powerful and 
elaborately controlled aileron system otherwise re¬ 
quired. 

It is interesting to note that the Germans, inde¬ 
pendently, seem to have come to a similar conclu¬ 
sion; their steep-angle guided bomb, Fritz X, was 
equipped with cross wings, originally at an obtuse 
dihedral angle, later at 90 degrees to each other. In 
steep-climbing antiaircraft rockets, the drag of the 
extra wing may be more serious. The Germans had 
not settled this question; of their four foremost anti¬ 
aircraft and two air-to-air missile projects, Rhein- 
tochter, Wasserfall and X-4 had cruciform wing sys¬ 
tems while Enzian and the Schmetterling missiles 
were banking monoplane structures. It is interesting 
to note that the former were under development by 
ordnance experts, the latter by airplane manufac¬ 
turers. 


14 21 Roll Stabilization 

The problem of wing configuration around the pro¬ 
jectile axis is intimately linked with the problem of 
roll stabilization. With cross wings or their equiva¬ 
lent it is possible, though not necessary, to stabilize 
the missile in some particular roll attitude relation. 
Such stabilization is advantageous where it is desired 
to maintain a certain geometrical correlation between 
a remote control gear on the one hand and range and 
line response of the missile on the other. 

This was the case of Roc 00-1000V a when con¬ 
trolled towards a collinearity program by direct 
vision in the Carp sight; a suitable type of roll stabili¬ 
zation gyro system, already developed for Azon and 
its descendants, was therefore adopted for Roc 
00-1000. Other types of roll stabilization, however, 
can also be adapted to Roc or the Roc missile to 
them, provided the guidance system is emancipated 
from geometrical correlation with the original target 
orientation and is based either on a televised image 
or is wholly automatic and bird-contained. 

One such roll stabilization system may be evolved 
from a gyroscope aggregate arranged to maintain the 
elevator axis horizontal. This has a certain advantage 
when it comes to compensating automatically for the 
computed gravity effect on path curvature, because 
then only the elevator system need be equipped with 
the gravity compensator; the rudder remains entirely 
unconcerned with gravity. This type of roll control 
may be advantageous when guidance is wholly by 
automatic target seekers or by television, not by 
direct vision or any other remote-control system 
which would lack recognition of vehicle orientation. 
In all other roll stability or control schemes, gravity 
will generally deliver a component into each of two 
orthogonal control planes, which may be aerodynam- 
ically equivalent (as with cruciform and annular 
wing systems) or which may perform different pitch 
and yaw functions (as in a truly banking vehicle). 

Where the gravity components are not to be com¬ 
puted but are expected to take care of themselves 
because the missile's control system is designed to 
sense higher derivatives of path errors, roll control 
can be relaxed. In fact the vehicle may be permitted 
to roll without preference for any particular attitude, 


a Roc 00-1000 is the designation of the Roc missile with ring 
wing and ring tail and 1,000-lb warhead; “-V” denotes the 
version adapted for direct vision guidance, “-T” denotes the 
version guided by television; Roc-X was the earlier experi¬ 
mental model with cruciform wing and empennage. 





268 


PROJECT ROC IN RETROSPECT 


but prevented from rolling too fast and making its 
intelligence “dizzy.” This system, in the opinion of 
many, exacts lesser demands of the gyro-dynamic 
apparatus than that of positive roll attitude control. 
Roll attitude remembrance, however, is necessary for 
a commutator to distribute the proper instantaneous 
pitch and yaw components where guidance is by 
remote control, or where gravity correction 15 is to be 
properly accomplished (without recourse to a tele¬ 
vised horizon 4 ). In the German antitank missile X-7, 
Max Kramer went so far as to make a monoplane 
vehicle deliberately roll at about 2 turns per second 
while flying on an essentially horizontal path; thus, 
antigravity lift was made only half the time, and a 
very fast response of the aerodynamic controls to the 
commutated signals was required. For this purpose 
he advocated the flow spoiler type of “interrupter” 
controls, which absorb remarkably little power. It 
would appear that this device conjures up a number 
of new difficulties for any precision intelligence sys¬ 
tem and for the dynamic response speed of the entire 
missile. Nevertheless, the fact that the design trend 
toward deliberate rolling of the missile was adopted 
in at least two groups of German projects® toward 
the later phase of the war, even though these projects 
were sponsored by individuals of considerably dif¬ 
ferent background, is notable. 

There is also another possibility of providing effec¬ 
tive roll control without attitude control of the body 
of the missile, namely, by articulating the wing sys¬ 
tem on the body so that the one can roll with respect 
to the other, stabilizing the wing system only but 
letting the body assume whatever roll attitude it 
may. This system has merits where the wing or lift 
controlling system is very compact (more so than on 
Roc) and of relatively low moment of inertia in roll. 
No such device seems to have been developed by the 
Germans. 

The free-flying missile has six degrees of freedom, 
which may be defined as three components of trans¬ 
lation and three components of rotation. When 
referred to an airborne system, the three rotation 
components are not of equal significance for the 
guidance of the missile; rolling rotates only the 
vector of aerodynamic force generated by the missile, 
whereas any angular motion in pitch immediately 


b In the early experimental Roc-X missiles, primary gravity 
correction of flight path was not attempted; hence no position 
gyro and no commutator were provided. 

°X--3, X-4, X-5, X-6, X-7, Rheintochter-3. 


produces changes in the magnitude of the force and 
thus affects the path curvature in its own plane 
directly. In yaw the situation corresponds to that in 
pitch if the wing is not a monoplane structure but 
cruciform, annular, or otherwise effectively symmet¬ 
rical around the longitudinal body axis. 


14 2 2 Lift Control 

If the (main) wing is entirely rigid and fixed to the 
whole vehicle, then its lift can be varied by merely 
controlling the angle of attack of the vehicle; this is 
usually done by means of a separate aerodynamic 
control element (elevator) at some leverage from the 
center of gravity producing a trimming moment. In 
order to achieve a stable motion under control, the 
wing and empennage system must be so controlled 
that a restoring moment is evoked by the accrual of 
an angle of attack until moment equilibrium is at¬ 
tained against the deliberately produced trimming- 
moment of the elevator. However, the tilting of the 
vehicle from the flight path takes time after the 
elevator has been applied; inertia comes into play, 
and a pitch oscillation is initiated, the dynamic 
stability of which requires investigation. The heavier 
the wing loading of the vehicle in comparison to the 
velocity head, and the shorter the coupling, the more 
serious this dynamic problem is likely to become. 
The leverage of the control surfaces may be positive 
or negative, corresponding to a Canard type of head 
controls on the one hand and to conventional air¬ 
plane tail controls on the other. In the former the 
control-surface lift is additive to the wing lift; in the 
latter, subtractive. 

In high-speed missiles the tendency is for the wing¬ 
span to be limited and not large in terms of fuselage 
width, as compared to conventional manned aircraft. 
Hence, in such a missile the lift generated by the 
body or fuselage when riding at an angle of attack is 
appreciable, like the dynamic hull lift of a dirigible 
airship. It is therefore possible to equip such a missile 
with a tail-less wing system and yet provide a fair 
leverage for control. Since an elongated body or fuse¬ 
lage is inherently unstable in attack, these tail-less 
wings must be positioned aft of the center of gravity 
and perform the duties of fins as well as wings 
(“fings”). The speed of response in attaining the at¬ 
tack attitude commanded and the dynamic damping 
of the ensuing pitching motion here become problems 
of special significance. 






WING SYSTEM 


269 


In Roc the original plan was to forgo the lift con¬ 
tribution of the fuselage or body and generate the lift 
exclusively by the main wing system as close to the 
center of gravity as possible, and to provide sufficient 
arrow stability to force the vehicle to ride at zero 
angle of attack of the fuselage. It was believed that 
this would simplify the control of self-homing and 
television guidance to such a degree as to justify the 
moderate sacrifice in maneuverability otherwise at¬ 
tainable. Encouraging results were obtained with 
experimental Roc-X missiles which had cross wings, 
radial interdigitated fins, and generous span at mod¬ 
erate wing loading. However, when it came to de¬ 
signing the heavier production article, clearance- 
space limitations for it underneath the wing of the 
Navy’s carrier airplanes led the design compellingly 
to the adoption of an annular wing as a more com¬ 
pact lift-generating device; because larger moments 
of inertia resulted, an equivalent dynamic stability 
was not retained. 

14 2 3 Trim Stability 

The trim-stability characteristics of Roc 00-1000 
are such that an initial trim disturbance would, ac¬ 
cording to calculations and extrapolations from wind- 
tunnel tests decay in the form of a weakly damped 
oscillation, the frequency of which may vary from 
about 0.25 c at low airspeed and density, to about 
0.75 c at high airspeed and density, while the ampli¬ 
tude would subside to about 60 per cent per cycle. In 
actual flight drop tests, oscillations of approximately 
the computed frequency have been observed but the 
story of their subsidence is not so simple. Occasion¬ 
ally the oscillations seem to decay slowly; on other 
occasions, however, they seem to persist, or even to 
increase temporarily. Whether these anomalies were 
the result of casual disturbances or a symptom of 
conditional or quasi-instability has not been defi¬ 
nitely established. It is possible that trim stability is 
dependent upon and sensitive to localized down- 
wash, wake, flow detachment, and interference ef¬ 
fects which may vary intricately with wing incidence 
in pitch and yaw as well as with Reynolds and Mach 
number, especially because of the flow constraint 
between the wing ring and the fuselage. The influ¬ 
ences may be more pronounced in the regime of 
higher speeds than those explored in the wind-tunnel 
tests that were made. It is also possible, in self¬ 
homing or television guidance, to fan such oscilla¬ 
tions by involuntarily introducing a phase lag be¬ 


tween oscillation and control. Any attempt at avoid¬ 
ing such fanning by smoothing out the control, or the 
missile’s response to it, over the period of a cycle or 
more is tantamount to a delay in control which may 
sacrifice precision of interception. On the other hand, 
it should be possible to damp the oscillation greatly 
and speed up its frequency by anticipation. A suit¬ 
able device may take the form of an autopilot boost 
system responsive to gyroscopically (or optically) 
determined pitch rotation, or better yet the form of 
an automatic elevator servo boost responsive to the 
local angle of attack at the tail, in order to increase 
both the decrement and the frequency of the oscilla¬ 
tion^ The booster must then be adapted to actuate 
an elevator at reasonable long leverage to the center 
of gravity. 

In Roc 00-1000 the entire wing ring is articulated 
in a Cardan joint. The moment of inertia of the wing 
system constitutes about 3 per cent of that of the 
entire missile; hence, according to the law of con¬ 
servation of angular momentum, the execution of a 
wing incidence command of 0-13 degrees would, 
aside from all aerodynamical trim effects, tend to 
produce an angle of attack of —0.4 degree. This may 
not be serious but theoretically it means that, to a 
small extent at least, incidence control will also 
superimpose a slight attack disturbance. 


14 2 4 Incidence Control 

With an annular wing, tilting the rigid ring as a 
whole on a universal joint is obviously the mechani¬ 
cally simplest method of incidence control, whereas 
with radial wings many other methods besides tilting 
the whole wing have merits; two have been consid¬ 
ered and tried on experimental Roc-X models—full- 
span flaps with internal and external balance. The 
choice between these and other possible incidence- 
control methods is mainly governed by aerodynamic 
considerations: (1) the lift coefficient should be a 
smooth, nearly linear function of the control dis¬ 
placement in the entire speed range; (2) the torque 
required to produce the displacement should be low 
in the entire speed range; (3) whether the torque is a 
monotonic function of the displacement or not is not 
necessarily of much importance so long as the control 
system is irreversible or gustproof by virtue either of 

d Such a booster will be even more important on guided 
antiaircraft missiles which have to be precisely controlled in 
much thinner air than guided bombs. 






270 


PROJECT ROC IN RETROSPECT 


friction or of the servo circuit chosen. A small 
unstable torque at low deflection is easily tolerable if 
it helps to keep the maximum restoring torque at 
high deflection a bit lower, provided the system is 
free from backlash. 

In the development of Roc, the design of the con¬ 
trol balance was based essentially on computation. 
Attempts were made in the wind tunnel to obtain 
rough checks of the adequacy of the servomotor 
drive to provide the necessary torque, but these ef¬ 
forts were seriously handicapped by the relatively 
low speed of 185 mph available in the wind tunnel 
and the uncertainty regarding the behavior of Reyn¬ 
olds number and Mach number effects at speeds 
higher than the test speed. The flight drop tests 
made with the finished Roc model revealed much 
useful information but they were not sufficient in 
number to afford opportunities to explore all the 
factors governing control efficiency and adequacy. 
Therefore, thorough wind-tunnel tests at full scale 
• and full speed and numerous systematic drop tests 
with complete instrumentation cannot be too highly 
recommended where scientific procedure is not over¬ 
ruled by the pressure of military exigency. 

14,2 - 5 Maximum Lift Demands 

The maximum amount of lift to be demanded of a 
missile has been the subject of a good deal of contro¬ 
versy. For a given design the maximum lift that can 
be generated is determined by the velocity head, the 
wing loading, and the stalling characteristics of the 
wing configuration; the stalling characteristics, how^- 
ever, may themselves be influenced by the Mach 
number. The maximum lift which can be borne by 
the structure without failure in strength or velocity 
may limit the permissible incidence at high velocity; 
brakes may be installed to safeguard against exces¬ 
sive loads. The highest maneuverability demands of 
any guided missile are likely to occur during the last 
phase of the flight in an effort to convert a miss into 
a hit. In a bomb (contrary to an antiaircraft projec¬ 
tile) this phase lies in the region of highest velocity 
head and hence highest maneuverability. Because 
both lift and centrifugal force evoked in a deflected 
path are proportional to the square of the velocity, 
the geometric effectiveness of an aerodynamic lifting 
device can be expressed independently of velocity F, 
in terms of path curvature 1/r forced upon a missile 
of mass M by a w ing system of effective w ing area A 
and at an incidence producing a lift coefficient C L 


(in the absence of a gravity component—for in¬ 
stance, as deflection from the vertical); that is, 


hence 


MV 2 

r 


h P AV 2 C L 



but it remains proportional to the air density p. 
Roc 00-1000 was designed to have a performance 
equivalent to a path-curvature radius of 7,500 ft 
from the vertical at sea-level density. At a terminal 
velocity of about 700 ft per sec, the centrifugal force 
evoked by such a turn amounts to approximately 2 g. 
Computations of sample trajectories had shown this 
amount of agility necessary to execute a sail-dive 
program according to the Carp technique to attain 
early collinearity at release speeds of the order of 200 
mph and altitudes of 15,000 feet and still leave a fair 
margin for the superimposition of manual remote 
corrective control. 

With television control, however, there is no com¬ 
pelling reason upon which to base a prediction of 
how much maneuverability will be necessary; such 
factors as field of vision encompassed, orientation of 
optical axis, wind or target speed to be overcome, 
and distance from which interception maneuver is to 
be begun, all have a bearing upon the maneuverabil¬ 
ity required to score a hit. To accomplish circular 
interception with controls full-over, the minimum 
distance at which leading must start is R 0 = 2 rU/V] 
hence for r = 7,500 ft and F = 600 ft per sec, 

U — 20 40 60 ft per sec 

Ro = 500 1,000 1,500 ft 

aside from any gravity component wdiich may be im¬ 
peding in a rear- or up-wind attack, or aiding in a 
head-on or dow r n-w T ind attack. 


14 2 6 Lift Control Rate 

The following questions now arise. How fast must 
the controls be capable of changing lift, say from 
zero to full-over, and should this rate be fixed or con¬ 
trollable? If controllable, then should the integration 
of the control speed, i.e., the determination of the 
control angle, be accomplished on the guider’s sta¬ 
tion or on the bird? This latter question asks whether 
it is the control speed or the wing incidence angle 
that should be signalled. In the experimental Roc-X 
missiles, control speed was signalled because the bird 



ROC GUIDANCE 


271 


was permitted to roll and no means were provided to 
ascertain the instantaneous roll attitude; hence, none 
were available to compute the proper incidence 
angles to allow for gravity. 

These points have been made: (1) the actual physi¬ 
cal quantity which it is desired to control in order to 
achieve interception is the position of the missile in 
space; (2) velocity components, acceleration com¬ 
ponents, and jolt components are the first, second, 
and third derivatives of position; (3) the higher the 
order of the derivative controlled, the greater the 
delay in executing the maneuver and in relaying the 
information of its success to the guiding agency. 
Hence, it should be desirable to guide by operating 
on as low a derivative as possible. Rudder or wing 
incidence essentially defines the transverse-path ac¬ 
celeration of the vehicle (second derivative of posi¬ 
tion), rudder speed the jolt (or third derivative of 
position). Deliberate control of the former thus ap¬ 
pears faster but it requires knowledge of other 
accelerations (for instance the gravity components) 
present. There is no doubt that incidence-angle con¬ 
trol is mechanically simpler, more foolproof, and in¬ 
trinsically more efficient from a power-conservation 
viewpoint than incidence-speed control. 

As to the speed attainable in moving the control 
element, it depends upon the servo power available 
and on the degree of balance of aerodynamic forces 
achieved; Roc’s control speed was not so fast as to 
make the inertia of the controlled mechanisms a sig¬ 
nificant factor. On vehicles controlled from the tail, 
secondary influences, such as downwash wakes and 
aerodynamic interferences, make it difficult to 
achieve high-precision balance over a large range of 
incidences with simple mechanisms (except possibly 
with spoiler types of control). Obviously the torque 
balance problem is more severe when the main wing, 
which generates the whole lift, is to be controlled 
than when only an elevator is controlled, which gen¬ 
erates only the small forces needed to serve as an air- 
impelled servo device and to operate on the main 
wing via the fuselage leverage. The advantage of 
direct wing-incidence control which is bought with 
the more precarious torque balance or power penalty 
should be a greater response precision of the vehicle 
as a whole, provided it is thereby relieved of disturb¬ 
ances due to angle of attack. Therefore, to fully 
justify the type of wing control of Roc, it would be 
desirable to complete the investigation into the 
origin of observed attack oscillations and into the 
prospects of devising means to prevent them. 


14,2-7 German Lift-Control Systems 

Some of the German designers have favored 
‘Trembling” spoiler controls, in which the amplitude 
of the control movements are almost always in excess 
of those needed and the speed of the control move¬ 
ment is also much faster than any response desired, 
the missile smoothing out the response by the effect 
of its own inertia in the course of integration over 
sensible periods. Even if the trembling is done at 
natural resonance frequencies, it would seem that 
this artifice is predicated upon precise balance and 
upon indirect control via elevator and fuselage angle 
of attack. 

The Germans also devoted a great deal of study to 
the theory of the so-called Black-White or Yes-No 
type of key control, in which the signals are not 
quantitatively modulated but pulsed, the length of 
each pulse governing the accumulation of response at 
a definite servo rate. The mathematical treatment of 
Yes-No types of control, especially when lags are 
taken into account, becomes rather involved because 
linearizing simplifications become treacherous. The 
advantages of intermittent control signalling systems 
presumably lie in greater simplicity and jamproof- 
ness of the receiver decoder for radio remote control. 
It would seem, however, that their application should 
be limited to devices whose tremble frequencies are a 
good multiple of the natural response frequency and 
whose torque requirements are not critical. 

14 3 ROC GUIDANCE 

14,3,1 Roc Radar Beam Rider 

As to the type of intelligence to which Roc was to 
be adapted, there have been repeated revisions of 
opinion during the course of the development. Roc 
started out as a vehicle for a radar beam rider. Later 
it was intended as a radar-homing bomb. Subse¬ 
quently, other homing target seekers were consid¬ 
ered, among them acoustic and heat-homing; photo¬ 
electric target seekers were actually tried out in drop 
tests. In 1944 a serious effort was made to press Roc 
into service for direct collinearity vision radio con¬ 
trol. Simultaneously it was developed as a vehicle for 
television control. These changes of policy were 
dictated by the advancement and delays in the de¬ 
velopment of the intelligence devices and by encour¬ 
aging or discouraging results obtained with them. 
Varying strategic considerations also bore an influ- 



272 


PROJECT ROC IN RETROSPECT 


ence on the program as World War II progressed. 

The beam-riding bomb project was shelved in 1943 
because it was believed (1) that it would be difficult 
to train the beam from the parent airplane on the 
target and to lock it thereon so precisely and steadily 
that the bomb would not be constantly subjected to 
excessive and abrupt control commands by the beam 
jitter; (2) the precise transmission of roll-phase in¬ 
formation down a slanting beam from a parent air¬ 
plane executing evasive maneuvers was deemed a 
difficult task at the time; (3) the accelerations im¬ 
posed upon the bomb when the parent airplane exe¬ 
cutes evasive maneuvers was appreciable, and con¬ 
siderable maneuverability would be required of the 
vehicle if it was to meet this contingency. It was then 
thought that the development of a radar target 
seeker responsive to an echo from a target illumi¬ 
nated from the parent airplane or another airplane 
farther behind would be a less ambitious task. 

Later experience indicated that this conclusion 
may have been fallacious. Although success was 
achieved in feeding the output of a radar-beacon¬ 
homing receiver into a differentiating and smoothing 
amplifier suitable to actuate the controls of Roc-X, 
flight tests aimed at measuring the reflection from 
foreign-illuminated target ships at sea gave discour¬ 
aging results in that reception from the upper quar¬ 
ters from which a moderately-steep-angle bomb 
would approach was handicapped by sea return clut¬ 
ter. Until a new way could be found by the radar 
experts to get around this difficulty, radar homing 
seemed limited to flying or gliding vehicles coming in 
on a flat approach path. 

Heat Homing 

While no attempt has been made to equip Roc 
with heat-homing devices for drop tests, it would 
seem to be reasonably certain that Roc should be at 
least as suitable a vehicle for heat homing as the Gulf 
bomb is in project Felix if there is any tactical advan¬ 
tage in the greater inherent maneuverability of Roc. 

Remote Guidance 

Any kind of automatic target seeker must com¬ 
prise a discriminator element which recognizes the 
target as such and distinguishes it from its back¬ 
ground and environment. Camouflage, decoys, and 
other ruses complicate the technical problem of 
target discrimination. It is therefore desirable to 


leave this part of the job to the human brain which 
still can apply judgment in a more versatile and 
delicate manner than any man-made machine char¬ 
acterized by anticipation and standardization. To 
apply judgment on the spot means either a suicide 
pilot or a guider close at hand but at a (relatively) 
safe distance, transmitting his services by radio from 
the parent airplane (or from an advanced ground 
station) to the falling bomb. 

14 3 4 Guider Station Control 

From the viewpoint of economy of the expendable 
article, it is naturally attractive to dispense with all 
target-sensing apparatus on the bird and to rely 
entirely on the recognition and intelligence directly 
available at the guider station. Fundamentally, this 
amounts to tracking both the target and the missile 
from the guider station and guiding the missile in 
such a manner that the two tracking rays coincide 
either at or prior to impact. Theoretically, for this 
purpose it should suffice to treat the missile as a mass 
point, provided its attitude is a unique function of 
the control-command history. Since unaided and un¬ 
disturbed ocular vision from aircraft has a discrimi¬ 
nating power of approximately }/% mil, it is tempting 
to rely on direct vision of bomb and target from the 
parent airplane for aim. However, it would be over 
optimistic to expect that the potential hitting ac¬ 
curacy would be equal to the visual acuity in detect¬ 
ing an error in aim. To arrive at a more reasonable 
expectancy, it is necessary to assess the accuracy 
with which the aim can be held continuously for 
some time prior to impact. On the other hand, some 
improvement of resolving power can be attained by 
resorting to moderate optical magnification in the 
sighting device, provided the telescope is gyrostabil- 
ized against vibration. A prerequisite of any point¬ 
aiming scheme is that the parallax between target 
and missile, during the critical guiding period at 
least, is either quite small or else rather accurately 
computed and subject to derivative measurement in 
control. An encouraging amount of success has been 
achieved by some techniques of collinearity or three- 
point-alignment guidance. This principle, including 
the stabilized telescope, was exemplified in this 
country by the Azon and Razon projects and the 
Franklin Institute’s bombsight projects Crab and 
Carp, and in Germany by both Max Kramer’s X- 
bomb projects and H. Wagner’s Henschel glide- 
bomb projects. 




TELEVISION CONTROL 


273 


Carp Apparatus 

Roc was brought into this picture as a compromise. 
It promised to have enough maneuverability to at¬ 
tain target collinearity after a sail-and-dive Carp 
program without imposing serious zoom-and-slow- 
down restrictions upon the parent airplane. On the 
other hand Roc’s potential feature of riding at or 
near zero angle of attack of the fuselage was wasted 
on a remote-vision technique in which the missile was 
treated as a mass point only. In fact still more lift 
per wing diameter can be attained if the fuselage is 
permitted to assume an appreciable angle of attack; 
a Gulf type of bomb with a fixed ring wing and a con¬ 
trollable tail would accomplish this purpose in a 
simpler manner than Roc with its more expensive 
Cardan-jointed wing ring. However, the pending- 
development of Roc as a television vehicle seemed to 
promise a fairly maneuverable missile to become 
available for Carp technique sooner and as a by¬ 
product, so that this course of action was adopted. 

Actually the development of the optical, naviga¬ 
tional, and practical aspects of the Carp apparatus 
and technique turned out to be a much larger project 
than was anticipated by its proponents. To expedite 
its completion would have required a more generous 
assignment of test* missiles and research manpower, 
as well as a deliberate advance development of all 
component parts of the system (flare, Carp sight, 
radio link, roll stabilization, and bird mechanism) 
before attempting to test the complete article. That 
a good solution for the collinearity-guided bomb 
would be of considerable interest there can be no 
doubt, especially in view of the fact that in antiair¬ 
craft techniques there is a definite trend toward 
making full use of the collinearity-aiming principle. 
The Germans claim to have achieved considerable 
success with it in drop and glide bombs. However, 
the problem of stabilizing the bird in roll so as to re¬ 
tain correspondence or range-and-line correlations on 
the bird and on the guide station will be aggravated 
when the guide-station carrier is allowed to perform 
evasive action to outmaneuver antiaircraft defenses. 

14 4 TELEVISION CONTROL 

14,4,1 The Observer’s Task 

The advent of Mimo television equipment small 
enough to be accommodated on a vehicle such as Roc 
raised the problem of how to interpret the television 


picture and how to govern the control command to 
be given in response to it. The observer of the tele¬ 
vision screen has to perforin the duty of the dis¬ 
criminator; he chooses and holds a target that can be 
recognized on the screen by the human eye but does 
not necessarily have such outstanding features that 
fully automatic apparatus could follow it while its 
image appears to change in size and perspective. 
Otherwise the guider’s function is but a link in the 
chain of an automatic regulator loop, though he may, 
depending on his understanding, experience, and 
skill, also introduce corrections to allow for known 
deficiencies of the remote control and regulating 
circuits. 

Attempts have been made to steer unmanned air¬ 
craft and automatic flying missiles by radio remote 
control toward a target seen by a television camera 
carried on the missile and observed on a television 
receiver elsewhere. These experiments have taught 
that to direct the missile into the vicinity of the 
target is one thing—to score a hit is another and 
much more difficult task. The reasons for the diffi¬ 
culty are not obvious. In fact it has been argued by 
some that steering the vehicle by watching the tele¬ 
vision screen should be no more difficult than to fly 
an airplane while looking through a small windshield 
and gunsight; even the apparent enlargement of the 
target image upon approach should not complicate 
the task beyond that of the pilot of a dive or suicide 
bomber. 

It seems that there are several major differences, 
to wit: (1) the pilot in a manned aircraft has a much 
greater field of vision, whereby he receives a great 
deal of helpful information, mostly subconsciously, 
partly by perspective, partly by reference land¬ 
marks; (2) the pilot receives vitally important sec¬ 
ondary information from other than the visual sense 
regarding acceleration and rotation, which have an 
influence upon the coordination between what he 
sees and what controls he has to apply to effect any 
desired course changes; (3) the pilot undergoes elab¬ 
orate training in spot landing so that he can practice 
the whole maneuver, except the very last suicidal 
phase ; e (4) in conventional aircraft, lift is made in 
one direction, and it is necessary to roll and then to 
“unroll” at the proper orientation. The rolling 
maneuver can be readily learned in nearly horizontal 
flight and in nearly vertical dive, but it is much more 

e It is understood that a remarkable number of Japanese 
suicide attacks missed their targets, indicating that pilots 
were insufficiently trained in the final phase. 





274 


PROJECT ROC IN RETROSPECT 


difficult to learn in moderately steep phases. In Roc 
this difficulty is altered in that the pilot has to learn 
to coordinate pitch and yaw instead of elevator and 
aileron controls, but the gravity component influence 
remains confusing. 

That training of the operator of television remote 
control is important has also been recognized, and 
simulators have been built in which the trainee sees a 
target area photograph gradually enlarge and move 
when the conventional control stick is manipulated, 
more or less as he would expect the image in a real 
flight to behave, while an instructor may introduce 
various disturbances. When operating such a simu¬ 
lator one begins to appreciate the magnitude of the 
task of learning how to parry these elusive disturb¬ 
ances. 

14 4 2 Aiming Aids 

There are reasons to believe that the job of learn¬ 
ing the technique of steering to hit the target can be 
vastly facilitated by relieving the operator of as 
much of the estimating, anticipating, and parrying 
as possible, in exactly the same manner as a gunner 
is helped by a lead-computing gunsight and aided 
tracking. This can be accomplished by inserting a 
regulator in the reaction chain which leads from the 
operator’s perception of the chosen target as some 
spot on the television scope screen to the application 
of a quantitatively metered-out amount of aero¬ 
dynamic elevator and rudder control. In a conven¬ 
tional aircraft control system, the control stick pri¬ 
marily governs the rates of turn of the vehicle but it 
causes secondary displacements of the image by 
virtue of the banking necessary to veer and of the 
change of angle of attack accompanying any changes 
of lift. Wind and target evasive maneuvers introduce 
unknown disturbances which conjure up the com¬ 
plications of interception of moving targets in space. 
In any vehicle descending at a steep slope angle, like 
Roc, information about the amount of trajectory de¬ 
flection due to the gravity component normal to the 
path tangent is not well conveyed by the television 
image; therefore, it appears as an unknown disturb¬ 
ance. 

In this connection it is interesting to note that 
during Mimo-Roc tests observers were led to experi¬ 
ment with two-man control, each taking care of the 
aim in one component direction only. This was meant 
to be a temporary expedient while operations were 
still hampered by various troubles. The Germans 


also experimented with two-man operation, though 
in a different manner: in collinearity guidance of a 
glide bomb, one man would roughly aim a master 
telescope while the other would sight through a 
repeater telescope and operate the remote control of 
the missile. Otherwise the Germans went to great 
trouble to render one-man operation more convenient 
and effective by developing means for training in 
proper coordination. Matters of this kind, involving 
psychological and physiological problems, cannot be 
solved dogmatically by theoretical analysis alone. It 
is for their solution that the advent of simulators is of 
particular significance. 

Definite advantages are seen by some in the idea of 
providing the bomb guider with a control device that 
reduces his job from that of a flier piloting an aircraft 
to that of a gunner aiming a gun or turret. The 
underlying philosophy is that this is much more a 
gunner’s than a flier’s job and that it is neither nec¬ 
essary nor natural to fly an aircraft in an aerobatic 
maneuver. In the ideal device, all the estimating, 
extrapolating, and leading would be relegated to a 
regulator or computer so that the gunner’s task 
would be reduced to maneuvering a reticule or bead 
onto the chosen target image. 

Suitable steering equipment will then comprise a 
reticule or bead system which is directly or indirectly 
movable over the surface of the television receiver 
screen. As to the most convenient method of moving 
this reticule, there are two schools of thought. 

In one, manipulation is accomplished by means of 
a direct mechanism, after the manner of a panto¬ 
graph pointer or a telautograph tracing a pattern; 
this method is most directly borrowed from the idea 
of aiming a gun towards a target and is particularly 
convenient when the inertia of the apparatus is 
negligible and where jerkiness of manipulation is in¬ 
consequential or self-averaging. The linkage of the 
mechanism can be equipped with suitable electric 
pickups which feed into an electric computer. The 
latter computes the control values to be signalled to 
the missile via the radio link. 

The other idea is to motorize the movement of the 
reticule and to control it after the manner of a motor 
control, speeding it up or slowing it down to “go 
after the target” or intercept it in due time. The lat¬ 
ter method requires a little more indoctrination to 
master but it has the advantage of lending itself 
better to a mechanization of the pickup of the reti¬ 
cule speed. 

An optical beam projector can be made to serve 





TELEVISION CONTROL 


275 


the same purpose as a mechanical reticule holder. 
This projector can be mounted in a universal joint or 
it may comprise two coordinate-line projector ele¬ 
ments, throwing a luminous image of a reticule pat¬ 
tern upon the scope screen from the outside of the 
oscilloscope but within the hood. The light can be % of 
a color contrasting with the fluorescent target image. 

Eventually an electronic reticule image can be 
developed, for instance, by producing a couple of 
bright (or dark) spots on one vertical scan line. The 
vertical lead motion would move the spots along the 
line, and the horizontal motion would select an 
earlier or later scan line. The aiming point could be 
the virtual center between the two pip spots so that 
the target itself is not blotted out by them. 

14 4 3 Camera Installation 

At a relatively early stage of the Mimo-Roc, it was 
arbitrarily decided to mount the Mimo television 
camera (fixed coaxially in the Roc head) looking- 
forward. It was hoped that the advantages of flying 
at zero angle of attack would reduce the reasons for a 
squinting or nodding eye. It was realized that several 
disadvantages would have to be suffered as a conse¬ 
quence of this design choice: (1) navigational lead 
would have to be limited to half the field angle en¬ 
compassed by the camera, (2) the target would not 
come into view until after about the first third of the 
fall time had elapsed, and (3) any tendency of the 
vehicle to suffer disturbances in angle of attack 
would have to be effectively subdued. 

As to the first of these three penalties, it was felt 
that a reasonable compromise could still be achieved 
with a field angle of 20 degrees, at which a properly 
led target having 20 per cent of the missile speed 
comes close to the image edge. As to the second prob¬ 
lem, a removable down-view prism was developed 
which afforded a view of the target during the flight 
approach before release, so that target contrast and 
television reception could be checked directly. The 
first phase of the drop would then be blind for about 
12 seconds. In order to enhance the probability of the 
target then coming into view, an unorthodox bomb- 
run-approach technique was developed which tended 
to minimize the crab angle. Similarly a variety of 
methods of choosing the release angle can be consid¬ 
ered: 

1. One is defined by the ballistic trajectory of the 
passive missile i.e., with controls zero, and deter¬ 
mined in a conventional bombsight-tracking pro¬ 


cedure; this obviously affords a maximum of physical 
possibilities of correcting random bad aims. 

2. Another method utilizes a late drop, requiring 
some early diving results in a very steep or nearly 
vertical impact which may be tactically desirable 
against certain types of targets and which has the 
advantage of minimizing the path-deflecting effect 
of gravity, possibly to the point where the effect can 
be neglected; the tactical disadvantages of a late 
drop are belated arrival of the target proper in the 
field of television and greater exposure of the parent 
aircraft to antiaircraft defense of the target. 

3. A sail-dive program, after a sufficiently early re¬ 
lease, has the same advantages without the excessive 
exposure to antiaircraft, but it may increase the 
chance of the missile veering too far off laterally. 

4. An early release with a relatively flat glide 
toward the target is a favorable technique as far as 
early television of the target and minimized exposure 
are concerned, but it requires more careful treatment 
of the path-deflecting gravity component, extends 
the flight time, reduces the impact speed and angle, 
and encroaches upon the forward-range-correction 
tolerance. 

14,4,4 Compensation for Wind 

and Target Motion 

Obviously it would be strategically most desirable 
to provide the missile with the equipment necessary 
to accomplish all these tactics and to leave their 
choice to the individual mission crews. Sufficiently 
precise gravity correction would have to be provided 
for this purpose. The question of what should be done 
to take care of the influences of wind and target 
motion, which are functionally equivalent, has come 
in for a great deal of attention. 

If a missile is so controlled that it tends to center 
the target in the field of a fixed eye, then an approach 
develops into a pursuit curve, and the question arises 
of how far the missile will miss under typical condi¬ 
tions of target or wind speed not being compensated 
for by navigational interception. The bird will miss 
if it does not lead the target because, in a pursuit ap¬ 
proach at large speed ratio, the rate of turn required 
to keep the aim of the target is small at first but in¬ 
creases toward the end phase until it finally exceeds 
the agility of the missile. The redeeming feature is 
that while the missile can at best wind up in its 
tightest circle osculating the pursuit spiral, the total 
miss may not be so large as to be objectionable, nor 



276 


PROJECT ROC IN RETROSPECT 


even larger than the inevitable errors due to other 
causes, especially when the wind or target speed is 
moderate. A rough idea of the order of magnitude of 
the miss can be readily given on the slightly opti¬ 
mistic assumption that the link and servo system lag 
is negligible. 

Under this assumption the miss is defined as the 
difference between the distance Ut yet to be traveled 
bv the target at speed U and the deflection of the 



Figure 1 . Angular relations in range-vertical projec¬ 
tion of pursuit curve. 

path from the last tangent at which maximum avail¬ 
able control was attained, i.e., at the critical range 
R c , under the combined influence of the path-normal 
accelerations due to this control and due to gravity. 
(See Figure 1 .) The available control is characterized 
by the minimum turning radius r = Vy, so that the 
acceleration due to it is y V = V 2 /r ; the acceleration 
due to the effective gravity component is < 7 -sin 7 . 

Hence, as soon as the path-curvature demand 
( U cos 7 )/Vr exceeds the available maneuverability 
due to aerodynamic control, namely 1 /r, plus (or 
minus) that due to gravity, namely (^siny)/^ 2 — 
that is, after passing the critical range R c = ( UV cos 


y)/(F 2 r -j- g sin 7 )—a miss of magnitude 
R C U/V is incurred. 

The following table gives an idea of these critical 
ranges and misses for various values of target (wind) 
speed U and terminal path slopes y c . 


0 


Critical range R c in feet 

Miss in feet 


Degrees 

U = 20 

40 

60 

20 

40 

60 

Tail-wind 

75 

167 

333 

1,000 

18 

17 

25 

or bow 

30 

225 

450 

675 

11 

23 

34 

attack 

15 

285 

570 

855 

14 

29 

43 


0 

350 

700 

1,050 

18 

35 

53 

Head-wind 

-15 

410 

820 

1,230 

20 

41 

61 

or stern 

-30 

455 

910 

1,365 

23 

45 

68 

attack 

-45 

420 

840 

1,260 

30 

42 

63 


It would seem that a quasi-scopodromic approach 
without automatic navigational lead is probably 
adequate for target and wind speeds up to about 
40 mph. Some deliberate lead can perhaps be applied 
by the bombardier, aiming for the bow of a moving 
ship, or to the windward in case smoke indicated a 
strong wind in the lower strata. 

The Mimo-Roc drop tests were conducted without 
benefit of a lead-computing device, although the 
need for some such device has been recognized and 
emphasized. It had been recommended that under 
reasonably favorable weather conditions none might 
be needed, at least until the television and control 
functions of the missile had proved practicable. As it 
turned out, about half of the drops were hampered 
by television troubles and the rest by angle-of-attack 
oscillations of the bird, either of which would have 
interfered with any precise lead computation, which, 
so far as it aims at interception of target motion or 
wind, depends upon a measurement of a rate of 
change of the target bearing as seen from the missile. 
This rate can be picked up directly from the target 
image speed as seen on the television receiver screen, 
provided the optical axis of the television camera re¬ 
mains pointed in the direction of the missile’s air¬ 
speed vector. The most direct method of accomplish¬ 
ing this measurement consists in forcing the vehicle 
to ride at zero angle of attack, like an arrow. Whether 
Roc can be made to avoid unfavorable angles of at¬ 
tack, either by moving the center of gravity far 
enough forward or by auxiliary stabilizing controls 
automatically responsive to the least angle of attack 
and more effective in wiping out any disturbance 
than the fixed empennage now provided, remains to 
be explored. Theoretically this should be possible, 



























TELEVISION CONTROL 


277 


but the apparatus required may turn out to be 
burdensome. 

14,4,5 Compensation for Angle of Attack 

If angle-of-attack oscillations and appreciable trim 
angles are not eliminated from the vehicle, then there 
are several possibilities of compromise. 

1 . The control command, or its execution, is made 
so sluggish as to smooth out the pitch and yaw oscil¬ 
lations, either by making it impossible for the guider 
to follow them with a geared-down manual control 
element (like the knobs tried out on the last two 
Mimo-Roc drop tests made thus far), or by means of 
electrical-mechanical smoothing devices inserted in 
the chain between the guider’s manipulation and the 
command execution. Whatever artifice of this kind 
is resorted to, it will of necessity reduce the alertness 
of the response to the real target bearing derivative 
and introduce a time Constant of the order of the 
period of the attack oscillation which is to be 
averaged. 

2 . The Germans were very conscious of the angle 
of trim because the vehicles on which they tried tele¬ 
vision were controlled by more or less conventional 
elevator systems. They experimented with television 
cameras mounted in a swivel joint on a sailplane and 
tilted mechanically by means of a pitch wind vane. 
For guided missiles they were sold on the idea of con¬ 
trolling at least the pitch angle of the optical tele¬ 
vision axis by a prism or mirror system governed by 
a pitch wind vane. In yaw, however, they believed 
that a banking monoplane missile could be endowed 
with sufficient arrow stability to avoid adverse yaw 
oscillations. In a vehicle possessing aerodynamical 
axial symmetry, like Roc, alignment of the optical 
axis with the flight path tangent, independent of trim 
oscillations, would necessitate mounting the camera 
in a universal joint f permitting several degrees of 
movement in any direction, and providing a means 
to govern the nodding motion by what amounts to a 
vector wind vane. Instead of building a mechanical 
wind vane, which has inertia and tends to oscillate 
itself, aside from being disturbed by the airflow past 
the vehicle, a three- or four-orifice pitch-and-yaw 
meter may prove easier to develop as a means to 
govern the alignment of the optical axis, and it can 
be made more rugged and serviceable. 


f A pitch and yaw mirror or prism system would be very 
clumsy at best. 


3. If the camera were mounted in a Cardanic sus¬ 
pension on the vehicle, but so softly sprung like a 
seismometric mass that it will not appreciably par¬ 
ticipate in the oscillation, then it would of necessity 
also lag behind the rate of turn of the missile in its 
curved trajectory, possibly frustrating the advan¬ 
tages of such suspension. 

4. In Germany, the DVL (Deutsche Versuchsan- 
stalt fur Luftfahrt) developed a pendulum scheme 
for orienting an optical system (a photoelectric con¬ 
trast seeker, not a television camera) parallel to the 
flight-velocity vector. The scheme is based on the 
idea that the direction of the resultant airforce 
(vector sum of lift and drag) is a unique function of 
the angle of attack, presumably known from wind- 
tunnel tests. The optical system is mounted on a 
pendulum which is articulated on the missile by 
resting on a (three-dimensional) evolute cam body. 
It seems that some means to damp the evolute mo¬ 
tion will be required to make this system practical; 
its self-containedness and freedom from aerodynamic 
appendages are distinct advantages. 

5. A more indirect method of allowing for the angle 
of attack of the missile is also conceivable. It pro¬ 
vides a pitch-and-yaw meter on the vehicle and 
transmits its readings to the guider or computer sta¬ 
tion via television or radio. If this is done via tele¬ 
vision, then the observer is burdened with having to 
read or interpret two pairs of coordinates from the 
screen, that of the target and that indicating the air¬ 
speed vector. If it is done by a radio signal, though 
possibly within the television channel, the data can 
be picked up and digested by a computer without 
bothering the observer. 

14 4 6 Interception 

If a computer is to be provided at all, then it is 
obviously desirable to make it do as much of the job 
as has a significant influence upon the hitting accu¬ 
racy. The theory of interception by means of auto¬ 
matic target seekers has been studied by many 
investigators. There are various ways of effecting 
interception of a moving target. The theory must 
recognize and introduce the various influences which 
affect the giving of a control signal, the control ef¬ 
fected thereby, and the result of the ensuing maneu¬ 
ver showing up on the scope screen and inviting 
further correction of the signal. The geometry of the 
orientation of the seeker with respect to the flight 
path has a dominating effect upon the presentation 





278 


PROJECT ROC IN RETROSPECT 


of a television picture and hence upon the guider’s 
reaction. Various methods of changing, biasing, and 
regulating the orientation of the seeker’s axis with 
respect to the vehicle’s flight path have been pro¬ 
posed and tried by different investigators; they may 
introduce an artificial, mechanically controlled angle 
of squint between the axis of the seeker and the fuse¬ 
lage and may have to take account of whatever angle 
of trim of the fuselage axis against the flight path in 
air may remain unsuppressed; or else they may con¬ 
trol the orientation of the seeker in space by gyro¬ 
scopic means and measure the space orientation of 
the vehicle independently by similar means. 

As to the interception maneuver to be executed 
by remote guidance, several alternatives offer them¬ 
selves. 

1 . The least that can be done is to make the missile 
turn early enough so as to intercept before the turn¬ 
ing requirements exceed the agility of the missile. 

2. The geometrically most perspicuous proposition 
is to distribute the turning evenly over the entire 
flight path; this implies approach along a circular arc 
path. It is the path of least peak curvature and ac¬ 
complishes the greatest path deflection for a given 
maneuverability. In the horizontal plane this is ob¬ 
viously also the method of least aerodynamic effort; 
in the vertical plane no similar criterion exists be¬ 
cause gravity and density variation confuse the issue, 
the aerodynamic effort required being but the sum 
or difference of the force required to produce the 
geometric path curvature and of the variable gravity 
component already tending to produce some path 
curvature. 

3. If ample agility is available, then it may pay to 
do the turning early, so that less turning is required 
later, and the path soon settles down toward a 
straight-line collision course which may be repeated¬ 
ly or gradually corrected to frustrate evasive action 
by the target. 

The geometry of interception can be conveniently 
expressed in terms of the angular velocity of the line 
of sight in a coordinate system to be anchored in the 
flying medium, oriented in such a manner that the 
angles to the line of sight r and to the line of flight co 
are measured from the normal to the line of target 
motion of presumably constant speed U with re¬ 
spect to the medium, viz., 


(U sin t — V) sin { u — t) + U cos (co — t) 
R 


( 1 ) 


which, for small lead angles co — r and for large 


speed ratio V/U, can be approximated by the first 
order terms: 


-F(co-r) u 

R ^ R 


(la) 


where V is the speed of the missile and u = U cos r, 
the target speed component normal to the instan¬ 
taneous slant range R. 

The angular acceleration of the line of sight is then 


( VtR - VtR - VccR + VcoR - uR) , 9 x 
R 2 ( ' 


The range shrinks essentially at the rate 
R = —V 


which condenses the above into 


(2 T - <b)V 
R 


(2a) 


This means that the angular acceleration grows large, 
hyperbolically as the range R shrinks to nothing 
when 

— < 2 (3a) 


co = 2r makes r vanish and produces a constant rate 
of turn co. If the proportionality factor co/f of control 
is not 2 but K, such that 

co = Kt (3b) 

then the choice of K < 2 causes the turning to lag 
and grow late, whereas K > 2 causes the turning to 
overdo at first and to diminish forthwith. A very 
large K leads quickly into a straight collision course. 

To accomplish control according to equation (3a) 
or (3b) would require the measurement of the angu¬ 
lar velocity f of the line of sight against a reference 
direction fixed in the medium or in terrestrial space. 
A practical method known to measure f directly con¬ 
sists of gimballing the eye on the bird, locking it on 
the target, automatically in the case of a target 
seeker or via radio remote control in the case of tele¬ 
vision and measuring its rotation gyroscopically, for 
instance, by governing the rotation through precessor 
motors or otherwise influencing a free gyroscope car¬ 
ried on the eye gimbal system. 

After f is measured, the required angular velocity 
of the line of flight is thus determined in proportion. 
1 he amount of aerodynamic force to be generated in 
order to produce the rate of turn dd can then be com¬ 
puted (if all ambient factors are known). 










TELEVISION CONTROL 


279 


If the eye is not gimballed but fixed in the bird, as 
was the case in the Roc missiles, then angular veloc¬ 
ity of the line of sight is not directly measurable with 
reference to terrestrial space but only that of the 
lead angle between the line of sight to the target and 
the bird axis, which latter is represented by the cen¬ 
ter of the television screen. The lead angle X, is cor¬ 
related to the absolute sight-direction angle r by the 
identity 

T + X = CO + Oi 

where a is the angle of attack, as illustrate 
2. If the latter is forcibly kept zero then 


comes 


hence 


co = Kt = K6) — K\ 


If n is some value other than 2, then X does not tend 
to be constant, but 

\ »-2 Z = £L 

X n - 1 R R (5) 

as can be seen by inserting equation (4a) into equa¬ 
tion (2a). Consequently, since 

R = Ro -j - Rt — R 0 — Vt 


(4) 


x ->.(§•)' 

(6) 

in Figure 

and 



(4a) 


X + u/V 

Xo + u/V W 

(7) 

then be- 

but for n — 2, 

X = X 0 and 




X + u/V R 

Xo d - u/V Ro 

(7a) 


K 


K - 1 


X = nX 


(3c) 


By way of examples: 

K = 1 corresponds to n = oo, pursuit course 
K = 2 corresponds to n = 2, circular inter- 

. ception 

K = oo corresponds to w = 1, immediate turn 

into straight col¬ 
lision course 


which states that an incidental lead angle has to ap¬ 
proach the collision-course angle linearly as the range 
shrinks during the approach. When n > 2 the turn¬ 
ing has to be done earlier, and for very large n it 
degenerates into the collision course. For n between 
1 and 2 the turning has to be delayed, and the ap¬ 
proach degenerates toward pursuit. For very strong 
negative n the pursuit tends to become clinodromic, 
i.e., it tends to maintain any existing lead Xo un¬ 
changed. 


Any automatic control scheme based on equation 
(3c) directly would become unstable whenever K 
dropped below 1, as this would imply a negative value 
of n. This difficulty can be overcome by designing 
the automatic control regulator so as to take into ac¬ 
count the fact that any deliberate change in cb will 
entail an identical change in X, because such a change 
will not immediately affect f, which is cb — X. Hence, 
the change Acb necessary to establish compliance 
with equation (3c) from a condition of non-compli¬ 
ance (coi, Xi) for any chosen value of n is 


LIFT PATH CURVE 



Figure 2. Angular relations in Roc trajectory. 


whereupon the new image velocity X indeed becomes 

^ coi — Xi 
a — : 

n — 1 

This means the servomotor system of an automatic 
regulator should be arranged to tend to reduce any 
difference between the existent value of cbi and a 
presently computed value cb = K(u>\ — Xi). 


The question of how the image speed X can be 
picked up from the screen has already been touched 
upon. It can be done either by manipulating a reti¬ 
cule to coincide with the target image and measuring 
the derivative of its position or by motoring the reti¬ 
cule after the target image and measuring the motor 
speed directly, always provided no angle of attack a 
exists to enter into equation (4), or that such angle of 
attack is suitably determined and its rate of change 
subtracted from X. 











280 


PROJECT ROC IN RETROSPECT 


Assume, now, that by this procedure and by suit¬ 
able choice of the interception factor K or n, the de¬ 
sired rate of turn u is determined. The knowledge of 
6 ) itself does not, however, suffice to determine the 
amount of aerodynamic control as governed by the 
wing incidence angle a and the lift coefficient Cl cor¬ 
related with it. The latter is determined by the equa¬ 
tions 

Q 

%pA V~ m = g sin c op + o)pV (8) 

if c op is measured from the vertical, for a wing control 
component so oriented that its lift is in the pitch 
plane defined by the vertical and the flight path, 
while 

hpAV^ = u r v (8a) 

without the gravity effect for the wing control com¬ 
ponent in the yaw plane, i.e., normal to the former. 
Hence 

n - 2M (“ + sin <*p-g/V) 

Cl - -—pAV - (8b) 

(the parenthesis term in pftch only). 

If the elevator and rudder planes are not roll- 
stabilized with respect to gravity, then each of them 
will be affected by a component of g times the sine 
and cosine respectively of a roll angle suitably de¬ 
fined with respect to the horizon. 

Since aerodynamic action depends upon air den¬ 
sity and airspeed, the control sensitivity to X should 
vary with these parameters. In other words, as the 
bomb drops into denser air and gains speed, the con¬ 
trol commands should be attenuated, the X term in¬ 
versely proportional to pV. Some German authors 
have emphasized the desirability of such progressive 
attenuation. Against it, the opponents of the idea 
have adduced the argument that since some leeway 
broadens the choice of the interception factor n, it 
ought to suffice to select a fair value for the middle 
run, condone some weakness of response in the early 
stage of the fall, and reap the benefit of sharper con¬ 
trol toward the end of it. 

There is, however, a solution which evades part of 
the issue by operating on the centrifugal force so that 
no knowledge or estimate of the air density is nec¬ 
essary. This scheme consists in signaling not a defi¬ 
nite control-surface angle but the desired accelera- 
tional load factor n v in pitch and n v in yaw, viz., 
u p — o) p V/g -f- sin oj p and n v = uyV/g respectively. 


Two servomotors are arranged to move the controls 
until two accelerometers match the signaled values. 
The airspeed still enters the picture. So does the 
method of roll stabilization. 

As to the influence of gravity upon the missile’s 
path, it can readily be seen that to compensate for 
it alone the bird should develop a lift coefficient 



upon which the lift demand for curving the path 
03 /V must yet be superimposed. (W is the weight of 
the missile.) Equation (9) applies to that plane of the 
vehicle which contains the gravity vector. There are 
several schools of thought regarding how the variable 
values of p, V, co should be determined and entered 
into the control-regulator system. 

14 4 7 Mimo-Roc Computer 

One school of thought prefers to arrive at the an¬ 
swer by means entirely contained on the parent air¬ 
craft so that no expendable apparatus need be in¬ 
stalled on the missile for this purpose. The computer 
developed by Division 7, NDRC, for Mimo-Roc be¬ 
longs in this category. The scheme amounts to an in¬ 
tegration of the motion of the bird under an assumed 
or recorded history of the control commands given 
or not given. Such integration is afflicted with some 
uncertainty because of the accumulation of inevit¬ 
able inaccuracies of zero setting and command execu¬ 
tion. This inaccuracy, if it is of any consequence at 
all, can theoretically be reduced by telemetering the 
velocity head Y 2 pV 2 and attitude a> on the missile by 
means of an airspeed indicator and a free gyro. 
Whether a greater precision might actually be real¬ 
ized in this manner is doubtful in view of the diffi¬ 
culties attending any scheme of telemetering, whether 
it is accomplished by existing television or by a sepa¬ 
rate radio channel. 

It has also been proposed to take care of the grav¬ 
ity influence by a series of trial-and-error glide-angle 
changes in which a straight collision course would 
eventually be reached after several discrete steps. 
The Germans tested this principle on one of their 
flat-glide bombs. The success of the scheme (which 
has a correlate in surface navigation) depends upon 
the time available and the precision of the deter¬ 
mination of 03 and X. In dive bombs it is hampered by 
variations of air density and airspeed, and by the 
brevity of time. 




TELEVISION CONTROL 


281 


14 4 8 Computation on the Missile 

Another method requires some expendable devices 
on the bird but has certain merits, especially that of 
being entirely independent of the television and 
command-transmission system—in fact of the entire 
command and lead-computing system. The gravity 
connection is separately computed on the bird and 
applied there. If it is to be done during all phases of 
the flight, an airspeed indicator must be provided to 
measure the velocity head }/^pV 2 and a free gyro to 
furnish co and thus sin co. The quotient of these quan¬ 
tities is then fed into the bird’s amplifier or servo 
system so as to superimpose the value indicated by 
equation (9) upon whatever command arrives. 

It so happens that the speed of the missile tends 
toward the condition of equilibrium between the drag 
and the axial component of gravity, viz., 

W cos co = \ P A D V 2 (10) 

Substituting 2W/pV 2 from equation (9) into equa¬ 
tion (8) makes 

Cl„ = (^p) tan « (11) 

where A D is the equivalent drag area of the missile, 
which is approximately known (except for the in¬ 
duced drag, which depends on the lift coefficients). 

Equation (11) shows that the correction is essen¬ 
tially governed by the slope angle co. It tends toward 
zero near the vertical dive but would be large if the 
path were flat. Fortunately, in Mimo-Roc flat-flight 
path, phases are of no concern because the target is 
not yet in the field of view. It probably suffices to 
apply this correction for angles of co greater than 45 
degrees. 

The angle co can be measured by means of a bird- 
borne free gyro which is uncaged upon release. Un¬ 
fortunately, the standard Schwien gyro aggregate, as 
it is installed in the direct-vision-controlled Roc-V, 
is not adapted to measure co because its free rotor 
rotates as a wagon wheel with its axis in an unsuit¬ 
able orientation. 

The job can be done by means of a free gyro which 
has its spin axis vertical, a tan co potentiometer on 
the outer gimbal, and roll-stabilizing contacts on the 
inner gimbal. The potentiometer could be so con¬ 
nected to the circuit of the present elevator-actuator 
follow-up potentiometers that it would bias the ele¬ 
vator radio signal by just the desired amount and 
make the actuator find the position, which is defined 


as the sum of the control command received by the 
radio link and the gravity-component correction. In 
this event the wagon-wheel gyro rotor can be dis¬ 
pensed with, the new spinning top rotor replacing the 
former and thus accommodating the whole gyro sys¬ 
tem in the same standard Schwien gyro housing. The 
roll-rate check gyro has to be retained. Although it 
cannot easily be coupled mechanically to the free 
gyro, the coupling can be accomplished electrically. 
The action of this new aggregate would not differ 
much from the conventional system in the early 
phases or so long as the flight path did not become 
very steep. If the missile came within a few degrees 
of a vertical dive, slight pitch or yaw control deflec¬ 
tions might call for large roll adjustments to keep 
the elevator axis horizontal. This might overtax the 
ailerons and confuse the observer. Since in this con¬ 
dition the gravity correction would be very small 
anyhow, no harm is probably done by preventing 
rapid rolling through the roll-rate gyro. 

The possibility of coping with roll stabilization by 
means of the standard Schwien gyro unit and of solv¬ 
ing the gravity correction by means of a supplemental 
free gyro with vertical axis, transverse inner gimbal 
axis and fore-and-aft outer gimbal axis also presents 
itself. In this case the tan co potentiometer would 
be driven between the gyro frame and the gimbal 
frame, and a sinometer between the gimbal and the 
bird itself. The sinometer would directly furnish the 
sine and cosine of the roll angle <f> as the bird rolls 
with respect to the horizon when it is called upon to 
veer from the original heading, so that range and 
line control remain as originally oriented, while the 
elevator axis does not stay in a horizontal plane. A 
suitable circuit would then have to be devised to sup¬ 
ply tan co cos <f> to the elevator bias and tan co sin <f> 
to the rudder bias. 

14 5 Tests and Conclusions 

In view of the multiplicity of methods and devices 
and techniques of steering a guided missile via tele¬ 
vision toward a target, too much emphasis cannot be 
placed upon the desirability of trying them out in 
nondestructive tests which duplicate or simulate the 
essential theoretical and practical aspects involved. 
The Germans recognized this need quite some time 
ago and spared no efforts to develop mechanical- 
electrical simulating devices. In connection with the 
Mimo-Roc venture a similar effort was made in two 
directions: one with an all-electronic simulator for 



282 


PROJECT ROC IN RETROSPECT 


the study of the fundamental principles, the other 
with a test cart on which two-dimensional intercep¬ 
tion problems with lead-computing control could be 
run. The lessons learned on these devices were most 
instructive. The history of this effort began at an 
early stage of Roc’s development, when the missile 
was still meant to fly without definite roll-orientation 
control. Here the spontaneous change of orientation 
posed a complex simulator problem for the solution 
of which a rather ambitious project of a model mis¬ 
sile test range was at one time seriously considered. 
In the meantime the NDRC group at Columbia Uni¬ 
versity developed electronic simulators of guided 
missile trajectory geometry to the point where the 
misses of individual runs could be quantitatively 
demonstrated without reproducing the physical mo¬ 
tion of the missile by airborne carriage. In fact such 
a simulator was put to practical use in training oper¬ 
ators who were to participate in the drop tests of 
visually controlled Roc 00-1000-V using the Carp 
technique. It afforded valuable opportunities to con¬ 
vey some of the dynamic response characteristic of 
the operation. That the Roc-V tests were neverthe¬ 
less hampered by practical difficulties of guiding 
should not reflect on the potential value of such simu¬ 
lators as training devices; rather it was due to the 
omission from the simulator system of the disturbing 
factors. This brings to mind that a simulator to be 
universally useful in the experimental stage of such 
a project must be versatile and equipped to add fac¬ 
tors and influences that turn out to be more signifi¬ 
cant than was anticipated. 

While the Columbia University group was working 
toward the development of more universal electronic 
simulators and of simulators adapted to the repre¬ 
sentation of television guidance, the Roc group un¬ 
dertook the construction of a mechanical model range 
suitable to simulate the most urgent problems of tele¬ 
vision guidance. In this model range the missile was 
represented by a steerable cart, moving at a speed of 
approximately 1:100 scale over a smooth floor to¬ 
ward a target independently movable on a cable 
track. Motion in only one plane was simulated, and 
roll was not represented, but devices were provided 
to approximate the path deflection by gravity when 
desired. Experiments were made with the real tele¬ 
vision camera mounted on the remotely steered cart. 
The bulk of the investigations into various guidance 
techniques was made without television, the guider 
(and an observer) riding the cart and observing the 
target space through equivalent optics. It is interest¬ 


ing to note that a similar training cart was developed 
by the Germans. 

No provision has as yet been made on the cart to 
simulate trim oscillation, although this was seriously 
considered. The cart was therefore applied mostly to 
the clarification of practical aspects of guiding with 
and without lead-computing aids. Various more or 
less primitive tracking aids or schemes were cursorily 
tried on the cart for educational purposes, but one 
scheme, for which the theory had been advanced un¬ 
der a project of Division 7, NDRC, was investigated 
more thoroughly. This scheme comprised an elec¬ 
tronic computer which would allow for target motion 
and gravity influence in one plane and aid the guider 
by means of a motorized reticule system. This device 
was built and operated on the cart under a variety 
of conditions. After some practice had been obtained, 
results were sufficiently convincing to encourage the 
construction of an airworthy prototype of the com¬ 
puter, which is available to be installed in the test 
airplane for future test drops. This procedure brought 
home the value of simulators of this kind as a means 
to iron out bugs or to dispel uncertainties regarding 
new and complex devices which have to perform as a 
link in a chain or regulatory loop. Appreciation of the 
simulator experience is not dimmed by the realiza¬ 
tion of the fact that some of the significant features 
were not portrayed in the model range, especially the 
behavior in trim and roll. 

It would be premature at the present stage of the 
cart studies to judge the relative merits of various 
computer schemes proposed or considered for appli¬ 
cation in conjunction with television. This much, 
however, can be safely concluded: the Division 7, 
NDRC, computer and reticule scheme does perform 
in any one plane as calculated and is far enough ad¬ 
vanced to be tried out in practical drop tests. Aside 
from this, it is felt that simulator and cart tests 
should by all means be continued and extended to 
cover various refinements and to afford training to 
those resuming flight-test operations. 

If an attempt is made to summarize the lessons 
learned in the pursuit of the Roc project, it must be 
borne in mind that this project was but a cog in the 
gearworks of a variegated guided-missile develop¬ 
ment directed by NDRC and not a rival of but rather 
a supplement to other missile projects. There seem 
to be enough different tactical merits in missiles trav¬ 
eling to their targets on differently inclined flight 
paths to warrant the development of specialized mis¬ 
siles. The definition of the border lines of their various 



TELEVISION CONTROL 


283 


applications may be somewhat arbitrary, but some 
overlapping among them can be accomplished by a 
judicious compromise between versatility and spe¬ 
cialization. 

The aerodynamics of missiles deserves to come in 
for more scientific investigation when higher speeds 
are to be attained. The supersonic regime, especially, 
is becoming a new field of expanding knowledge in 
which the pursuit of guided-missile projects is bound 
to be in the forefront and to have an influence upon 
the progress of the entire art, including the design of 
manned aircraft for high speeds. The choice of a suit¬ 
able aerodynamic configuration for guided missiles is 
dictated by considerations peculiar to them and in¬ 
separable from the guidance scheme and its tactical 
implications. This complex network constitutes a 
regulatory loop, the satisfactory solution of which 
can be achieved only by close coordination between 
the specialists in the fields of aerodynamics, regula¬ 
tion, and intelligence. Because of the importance of 
coordination between these fields, it is believed that 
the responsibility for the coordination should be 
placed at an industrial level upon a prime contractor 
for the entire system. 

The need for facilitating the dissemination of mu¬ 
tually pertinent data among the workers in this and 
associated fields has been impressed upon the per¬ 
sonnel working on the Roc project. Since the ending 


of hostilities and resulting relaxation of security 
measures, the exchange of information has brought 
to light some cases of duplication, but, more impor¬ 
tant, it has answered questions on unsolved problems 
and posed questions on important new problems. It 
is felt that a system which will allow the exchange of 
information in this field of endeavor under wartime 
restrictions must be developed. 

The flight testing of guided missiles is likely to be¬ 
come a discipline of ever increasing importance and 
scope. Here again the usefulness of close cooperation 
between those responsible for the conduct of the tests 
and the recovery of instructive information, those 
responsible for the design of the missile and its con¬ 
trols, and those who determine its tactical purposes 
and performance specification cannot be overempha¬ 
sized. In order to facilitate this cooperation during 
the flight testing of the missile, it is believed that 
adequate flight test facilities should be developed, 
with satisfactory housing and laboratories necessary 
for maintaining a high scientific standard in these 
intricate and significant tests. 

In closing, one is tempted to reaffirm the belief that 
the missile, its payload, propulsion, stabilization, 
controls, guiding, and launching and all the appara¬ 
tus necessary to make it operate, constitute an in¬ 
tegral system which must be considered and treated 
as a whole. 





Appendix A 

INTERCEPTION AND ESCAPE TECHNIQUES AT HIGH SPEED 

AND HIGH ALTITUDE 0 


During 1941, under contract from the U. S. Army Air 
Forces, Douglas Aircraft Company undertook a study to de¬ 
termine the effects of high speed at high altitudes on the prob¬ 
lem faced by fighters when attacking bombardment aircraft. 
The report which arose from that work comprises the body of 
Appendix A. The problem of maneuvering a fighter plane 
against a high-speed bomber so as to bring the fixed guns to 
bear is one of flying a pursuit course. The work of Douglas, 
therefore, involved a searching analysis of the dynamics of 
pursuit curves, and although the content of the contract in¬ 
volved no work on guided missiles, its outcome has become a 
fundamental classic, which has been invaluable to all the Divi¬ 
sion’s contractors concerned with the development of homing 
missiles. [Editor] 

The study which led to the essay, Interception and 
Escape Techniques at High Speed and High Altitude, 
was originally undertaken in 1941 in conjunction 
with a design study of a fast long-range bombard¬ 
ment airplane in an effort to arrive by theoretical 
analysis at some judgment of the merits of a design 
trend toward safety through high performance 
which was then beginning to take shape in the minds 
of visionary engineers. The idea was still a bit ahead 
of contemporary concepts of immediate necessities of 
the military situation, in which England was seen as 
hard pressed and in need of relief while America was 
not yet actively embroiled in the world conflict. The 
high-performance bomber project was shelved at the 
time, but the principle was eventually incorporated 
in the later development of the B-29 bomber. 

The report on interception and escape techniques 
was written at the time as a part of the project pro¬ 
posal and not intended to be published by itself or to 
pose as a comprehensive treatise on the subject. It is 
indeed but a very incomplete part of a survey of the 
field that could be covered or implied by the title. 

In retrospect, presentation of the report without a 
bibliography of the subject matter which has been 
a fruitful field for many investigators would seem 
somewhat presumptive. As a matter of fact, no time 
at all was allowed for a literature search in view of 
deadlines for the submittal of the project proposal. 

This fact also contributed to the decision of coining 
code words for technical terms which are peculiar to 
the science of pursuit, and for which more common 
words were in rather lax and ambiguous usage, though 
some of them (e.g., pursuit curve, dog curve, squint- 


a Written October 23, 1941. 


ing, vector sight, and many others) have since been 
adopted more generally, while others have entered 
the language of gunsight, radar, and antiaircraft fire- 
director techniques. The “Greek” words introduced 
in the present report were not proposed for universal 
adoption but merely to fix the ideas until coordina¬ 
tion with the nomenclature of related techniques 
could be accomplished. 

An admirable bibliographic survey of mathemati¬ 
cal treatments of pursuit problems has since been 
prepared for the Special Devices Division, Research 
Section, Bureau of Aeronautics, U. S. Navy Depart¬ 
ment, by Tufts College Mathematical Research 
Project. It is entitled Project RM-6, Mathematical 
Analysis of Ordinary and Deviated Pursuit Curves. 
A survey of some pertinent German papers produced 
during the war is expected to become available in the 
near future. 

In Sections A.3 and A.5 various conclusions were 
drawn regarding the tightest turns that could be 
flown in combat maneuvers. The analysis did not 
take into account the possible variability of the maxi¬ 
mum lift coefficient of wings with Mach number. In 
the light of knowledge and experience gained in the 
meantime regarding this variability, some of the con¬ 
clusions drawn may require revision. Whether the 
pursuer or the pursued is likely to be more hampered 
by premature stalling or flight stability difficulties 
will depend upon aerodynamical details of either 
craft. 

Section A.8, which deals with problems familiar 
to artillery experts but unfamiliar to the aeronautical 
engineer, is particularly sketchy. It was meant to be 
elaborated after consultation with experts in this 
field, but the report had to be concluded before the 
necessary information could be gathered. This sec¬ 
tion must therefore be considered as a mere collection 
of thoughts and an enumeration of problems rather 
than their solution. 

Section A.9 must also be forgiven as somewhat 
cursory in view of the vast amount of study which 
other agencies concerned with the training of pursuit 
pilots have actually devoted to the task of simulating, 
teaching, and practicing pursuit and interception 
flight techniques, even though the author did go to 
the trouble of verifying some of the statements made 
by improvised flight tests. 

Occupation with the thoretical aspects of air-to-air 
interception at the time the investigation was carried 


285 



286 


APPENDIX A 


out gave the organization a background which served 
it in good stead when it was called upon to contribute 
to the development of guided missiles, and particu¬ 
larly of Project Roc. 


ai SUMMARY 

The purpose of the following investigation is to 
anticipate and analyze the armament requirements 
for a radically new bomber project designed for such 
high speed and altitude that its own performance 
would render it extremely hard to intercept. The 
study envisages performances that may sound fan¬ 
tastic but they can be shown by proven aerodynamic 
methods to be entirely within the realm of immediate 
realization; they are probably no more in advance of 
the aircraft now in service than the latter are of those 
of but a few years ago. 

The results of the present study bear out the con¬ 
tention that the need for defensive armament of a 
high-altitude bomber decreases as its speed ap¬ 
proaches the value at which air compressibility in¬ 
fluences begin to impede further progress seriously. 

Limitations of Interception 

A raider entering enemy territory in the strato¬ 
sphere, safe from ground fire, can penetrate enemy 
territory to a considerable depth before he can be 
challenged by interceptor aircraft which have to be 
designed to match the raider’s high ceiling. The 
chances of being found by an interceptor depend on 
the latter’s climb and speed advantage, which can 
at best be small, and on an excellent, alert, detecting, 
warning, dispatching, and control organization on 
the ground. 

Interception vs Straight Escape 

Even the fastest interceptor has very little chance 
to pour effective fire into an almost equally fast 
bomber, except in a straight tail chase. Therefore, 
rear armament covering a moderate tail-cone field is 
all the armament definitely necessary on the bomber. 
High accelerations and unfavorable leading condi¬ 
tions hamper or thwart any close approaches from 
blunt angles. Unless he is far out of range, the pur¬ 
suer has to turn so fast to keep his bead on his quarry 
that the accompanying high accelerations would 
greatly impair his aiming accuracy, or even cause him 
to black out. A variety of aiming and approaching 
techniques are studied in detail. The practical limits 
depend on speed and effective firing range but it is 
shown that for a very fast, high-performance bomber 
an interceptor cannot practically bring fixed forward¬ 


firing guns to bear on the target closer than 500 yd 
from angles greater than about 30 degrees, or even 
closer than 1,000 yd from more than 60 degrees off 
the tail of the target without suffering more than Zg 
acceleration. 

Attacks from dead ahead appear to be extremely 
unlikely because of the terrific speed of approach, 
which leaves but a few seconds for recognition and 
decision, less than one second for fire at effective 
range, and but a fraction of a second to maneuver 
out of the way to avoid collision. 

The influence of properly leading the target is in¬ 
vestigated. It is shown that this refinement must be 
taken into account in order not to overestimate the 
acceleration suffered in the homing approach. How¬ 
ever, it is also revealed how complicated a job is the 
accurate determination of the factors governing the 
aim correction in any homing maneuver but the tail 
chase. These considerations may help justify the lim¬ 
itation of the bomber’s defense gunnery to a rear 
field cone. 

It is shown that cross-passage combat phases are 
theoretically possible at all angles without the inter¬ 
ceptor suffering acceleration loads. However, it is also 
shown how utterly brief such encounters would be, 
how few bullets could possibly sweep through the 
target, and how difficult it would be for the inter¬ 
ceptor, in the few seconds available, to determine and 
attain the best intersection course which would bring 
him into effective range just at the correct instant. 

The only way for the interceptor to challenge the 
high-speed, high-altitude bomber effectively from ap¬ 
preciable angles off the tail, and thus require him to 
defend himself there, is by mounting slant guns. 
Turrets would probably cost too much drag, reduce 
the speed advantage of the interceptor, and defeat 
their own purpose. Moreover, even pursuers with 
fixed oblique guns cannot very well attack effectively 
from very large angles because of their own maneu¬ 
vering problems and windage influences. 

Combat Maneuvering Techniques 

It is shown that the bomber, if equipped with ef¬ 
fective tail defense, can maneuver so as to force an 
attacker into a tail chase in which the bomber has an 
aerodynamical and tactical advantage. Also, he can, 
if he so elects, assume the initiative in a combat; it 
is shown under what conditions the bomber can frus¬ 
trate or delay an attack by certain dogfight ma¬ 
neuvers. 

Multiple Interception 

In a study of the possibility of multiple intercep¬ 
tion, the difficulties of accurate coordination of the 



LIMITATIONS OF INTERCEPTION 


287 


interceptors at high speed appear formidable so that 
on this score the bomber would not seem to require 
increased fire coverage in space. Simple escape ma¬ 
neuvers are likely to force simultaneous attackers 
into the same tail quarter to be resisted by the tail 
gunnery as they come into range. 

Mass Raid Techniques 

Consideration is given to mass formation flights 
and to the influence of mass formation tactics upon 
the armament requirements. While even in this case 
a tail defense is probably adequate, the tactical ad¬ 
vantage of equipping at least some of the bombers 
with forward-firing fighter guns is brought to light. 

Considerations of Effectiveness of Fire 

An investigation into the data required to establish 
some sort of probability calculus for the chances of 
receiving fire from definite quarters leads, although 
not to any rigorous calculation, at least to qualitative 
confirmation of the overwhelming predominance of 
the tail chase if the attacker is equipped with fixed 
forward-firing guns. The appearance of interceptors 
equipped with oblique guns pitched within reason¬ 
able limits would favor a proportional increase of the 
angular coverage of the tail cone. 

Mock Interception to Scale 

Some rules are established for the interpretation of 
practical studies of high-speed interception problems 
by means of reduced-speed mock-maneuver flight 
tests which may serve to practice various intercep¬ 
tion phases and to throw some light on their respec¬ 
tive practicality and seriousness. 


A2 LIMITATIONS OF INTERCEPTION 

The present study was prompted by an endeavor 
to find answers to the query: To what extent is a 
long-range, high-altitude, ultra-high-speed bomber 
safe from attack by conventional defense means, or 
what specialized defense means and technique would 
have to be devised to combat it? 

A 21 Bomber’s Speed and Altitude 

The advocates of ultra-high performance of the 
bomber as its best protection contend that armor and 
armament can be sacrificed or at least reduced to 
protection from very limited critical quarters for the 
benefit of speed and ceiling to the point where it can 
escape any effective defense against it that can be 


developed in the time before it becomes obsolete. To 
fix the idea, assume that the bomber approaches the 
enemy area in the stratosphere, say at 40,000-ft alti¬ 
tude, at a speed of V = 450 mph, so high and so fast 
that any attempt at ballistic ground defense would 
be ineffective and useless. As to air combat, the idea 
is to fly the bomber so fast that it approaches the 
speed limit set by compressibility burble drag so 
closely that the interceptor, who is of necessity sub¬ 
ject to similar limitations, is left but little speed ad¬ 
vantage with which to catch up with his prey before 
he runs out of fuel; his pursuit maneuvers are further 
handicapped by the high load factors which accom¬ 
pany any turns at high speed. 

A 2 2 No Standing Patrol 

If the defender were compelled to set up continu¬ 
ous aerial standing patrols of pursuit airplanes at high 
altitude, he would require a fleet of many thousands 
of airplanes constantly devoted to no other duty than 
this waiting job, a gigantic waste that might surpass 
in drain of resources the potential bomber damage it 
is set up to avert. One can try to estimate the size of 
such a curtain defense fleet, but for the time being 
the idea will be dismissed as fantastic. 

A 2 3 Defense Detector System Assumed 

Disregarding the possibility of the development of 
radically new weapons as yet unknown, we shall as¬ 
sume that both the bomber and the interceptor are 
equipped with directive radio detectors enabling 
them to spot and locate enemy aircraft in the air 
within a range of several miles (say 5 to 20), and that 
the defenders possess a highly organized scanning and 
warning system on the ground along their border by 
means of which they can detect and identify any in¬ 
vading bomber force while it is still at a distance D 
of many (say 30) miles away. The possibility of dis¬ 
abling the detectors by jamming (or electric camou¬ 
flage) may be admitted for close range in the air but 
it can probably be dismissed as far as the ground 
warning system is concerned. 

A 2 4 Interceptor’s Climb Required 

In order to be able to intercept the bomber before 
it reaches the shore or boundary of the defender, the 
time T h available and the average climbing speed C 
required from the instant of detection to the arrival 
at the invader’s flight level in the stratosphere (alti¬ 
tude H above the defender’s take-off field) are de¬ 
termined by the equations T H = D/V = H/C. If V 
is taken in mph, D in miles, and H in feet, then T H 
and C have to be multiplied by 60 in order to give 



288 


APPENDIX A 


the time in minutes and the average rate of climb in 
conventional units of feet per minute. Table 1 gives 
an idea of the magnitude of these quantities for 
V = 450 mph. These figures bring home to what de¬ 
gree the interceptor is handicapped by pressure of 
time, even if no time were lost in transmitting all 
necessary information and dispatching the pursuit 
formations. 


Table 1 



D (miles) 

15 

30 

45 

60 

H (ft) 

T h (min) 

2 

4 

6 

8 

30,000 

C (fpm) 

15,000 

7,500 

5,000 

3,750 

35,000 

C (fpm) 

17,500 

8,750 

5,833 

4,375 

40,000 

C (fpm) 

20,000 

10,000 

6,667 

5,000 


A 2 5 Undefendable Zone vs Climb Time 

As a reasonable example, if the detection distance 
is 30 miles, that T 0 = 7 minutes (time lost in trans¬ 
mitting the message and in take-off preparations), 
and that the actual climbing time to 40,000 ft will 
be T c = 20 minutes, there remains an undefendable 
zone of over 200 miles into which the bomber may 
penetrate unchallenged. For various other conditions, 
the depth D u of this undefendable zone, as deter¬ 
mined by the relation D u = V(T 0 + T c ) — D, is 
given in Table 2. The last column of Table 2 indi¬ 
cates to what extent the detection range will have 
to be increased to assure defense preparedness right 
up to the border. 


Table 2 


To 

Range of detection D 



Range of detection 

+ T C 


(miles) 




required for com¬ 

(min) 

60 

45 

30 

15 



plete defense (miles) 

20 

90 

105 

120 

135 

— 

s8 


150 

24 

120 

135 

150 

165 

•V 

- 


180 

28 

150 

165 

180 

195 

£ 

a> 

— 

210 

32 

180 

195 

210 

225 


o 

N 

240 

36 

210 

225 

240 

255 

c 

270 


A 26 Interception Maneuvers Required 

However, even after the interceptor arrives at the 
altitude of the bomber and finds it, he still has to 
maneuver into position for attack. The success of this 
maneuver will depend greatly on the conditions and 
technique of this approach. Consideration will be 
given to a variety of such conditions and techniques; 
although this survey may not be complete, it is in¬ 
tended to cover at least the critical phases. 


A 3 INTERCEPTION VS STRAIGHT ESCAPE 

In this part of the investigation it will be assumed 
that the bomber crew, because of their high speed, 
feel so secure and invulnerable that they fly serenely 
on toward their objective on a straight and level path 
without paying any attention to enemy efforts at 
challenging or intercepting them. This assumption 
will be dropped later in this appendix, but it is at 
once apparent that it is much to the advantage of 
the bomber if it can afford to minimize any depar¬ 
tures from its direct course towards its objective or 
at least toward a feint objective. 

A 31 Homing Pursuit (Scopodrome) 

Let it first be assumed that, after the interceptor 
has located and chosen his quarry, he pursues his 
target in a homing pursuit curve, aiming his airplane 
constantly at his victim without leading his targets 
The maneuver winds up asymptotically in a stern 
chase. The pursuit pilot holds his fire until he has 
arrived at close enough range and close enough on 
the tail to require but negligible target lead. 

This scopodromic concept of flying in the direction 
toward the target represents an idealized technique, 
introduced here for the sake of mathematical sim¬ 
plification, close enough to be fruitful but predicated 
upon several artificial assumptions. One of these is 
the supposition that the interceptor is equipped with 
a sight which tells its pilot the direction of its in¬ 
stantaneous flight path. In the past, airplanes were 
not so equipped. Their gunsights were bore-sighted 
with respect to their guns, which usually were in¬ 
stalled parallel to the normal speed flight path in 
straight flight. In a turn, however, the angle of attack 
is increased to overcome the centrifugal force, and 
this variation was not revealed by the conventional 
gunsight. An estimate of the influence of the angle of 
attack with a fixed gun and fixed gunsight will be 
made separately. 

Interceptor’s Speed Constant 

Assume that the interceptor approaches at constant 
airspeed v, neglecting the fact that as he turns his 

b Reasons for not leading the target may be that (1) at long 
range he may be uncertain as to the enemy’s flight direction 
and (2) later at moderate range because he may want to evade 
the complication of allowing for a variable lead vector as he 
banks in the approach turn while still beyond firing range. The 
term “homing” is not any too apt here, though derived from 
beam-flying technique with wind. It may be well to coin spe¬ 
cial words to identify the various approach techniques for 
ready reference. The term “scopodromic” is proposed for the 
present method; the word signifies “driving so as to move 
toward the visible target.” 

















INTERCEPTION VS STRAIGHT ESCAPE 


289 


drag must increase. (An estimate of this neglected 
influence will be introduced later.) 

Horizontal Plane Scopodrome 

Further assume that the pursuer either does not at¬ 
tempt to dive on his prey or that any advantage de¬ 
rived from such a dive can be expressed as a slight 
increase in his speed advantage over the bomber. 


Range vs Azimuth 


The polar equation of this horizontal scopodromic 
pursuit curve in terms of instantaneous range r and 
azimuth a of the pursuer's position off the pursued's 
tail can be written: 


r 


r T 


( tan f y 

sin a 


(i)° 


e = v/V, the speed ratio of pursuer to the pursued, 
which is here assumed slightly greater than unity, 
say 10:9 for a 500-mph interceptor or a 450-mph 
bomber; rr is the range at which the pursuer was 
headed at right angles to the flight path of his quarry. 
If he does not take up the chase until he is in a rear 
quadrant at an azimuth a 0 and at a range r 0 , then 
rr may be reconstructed by following the process in 
reverse, or else the same curve can be expressed in 
terms of any coordinated initial values of a 0 and r 0 , 
namely: 


r 

ro 


/ tan . 

I 2 1 sin ap 

I , ao I sin a 

\ tan 27 


( 2 ) 


As the pursuit approaches the tail chase for which 
a nears 0, equation (2) approaches its first order 
term, namely, 


r_ 

To 



( 3 ) 


Typical approach curves of this character are 
shown in Figure 1 in a coordinate system assumed 
traveling with the pursued aircraft for speed ratios 
of pursued to pursuer 1/e = 1.1, 1.0, 0.9, 0.8, 0.7, 
0.6, and 0.5. Of these, 1.0 is a limit case because, with 
speeds equal, the pursuer never catches up with the 
pursued; he can approach him no closer than 
The case of V :v = 1/e = 0.9 is a practical example. 


Accelerations 

The question now arises: What centrifugal “ac¬ 
celerations" and what load factors would be suffered 


0 For derivations see Section A.4.1. 


by the pursuer in such a scopodromic approach and 
what consequences will they have? The centrifugal 
acceleration in terms of a horizontal load factor n v 
is simply the product of the tangential and angular 
velocities of the pursuer; thus, n^g — — vda/dt. The 
angular velocity, however, is —da/dt — V sin a/r, 
so that 

vV 

n v g = — • sin a (4) 

This is constant for constant r /sin a = vV/n^g, which 
is the diameter of a Thales circle, tangent to the pur- 
sued’s path at the pursued’s instantaneous position. 
Such Thales circles are isobars , i.e., loci of equal load 
factors for the pursuer. A family of such circles, with 
the corresponding resulting load factors 

n = Vl + (5) 

annoted, are shown in Figure 2, together with scopo¬ 
dromic pursuit curves and range circles for the ex¬ 
amples v = 500 mph and V = 400 mph (1/e = 0.8) 
and V = 450 mph (1/e = 0.9). Table 3 shows at 
what ranges and azimuths certain load factors would 
be suffered in scopodromic approach. If the pursuer 
enters the innermost n = 5 circle, he would probably 
black out. 


Table 3. Range in yards (v = 500 mph) for various 
pursuer’s resultant load factors and azimuths. 



n 

2 

2.5 

3 

3.5 

4 

4.5 

5 

V 

a n v 

1.732 

2.29 

2.83 

3.35 

3.87 

4.39 

4.90 


90° 

2,895 

2,190 

1,773 

1,497 

1,295 

1,142 

1,023 


60° 

2,510 

1,897 

1,535 

1,297 

1,123 

989 

887 

1 

56°17' 

2,405 

1,820 

1,547 

1,472 

1,243 

1,077 

948 

851 

O 

45° 

2,045 

1,252 

1,058 

916 

807 

723 

lO 

30° 

1,447 

1,095 

887 

749 

648 

571 

512 


15° 

749 

567 

459 

388 

336 

296 

265 


90° 

2,570 

1,945 

1,575 

1,330 

1,152 

1,015 

910 

43 

60° 

2,225 

1,687 

1,365 

1,152 

998 

879 

788 

G. 

£ 

51T9' 

2,005 

1,518 

1,228 

1,038 

898 

792 

709 

8 

45° 

1,840 

1,375 

1,113 

940 

813 

718 

643 

30° 

1,285 

973 

788 

665 

576 

508 

455 


15° 

-666 

504 

408 

345 

299 

263 

236 


For any given resultant load factor n, the diameter 
of the Thales circle or cross-path range is: 


r T = 


vV 


\/n 2 — 1 


( 6 ) 


The banking angle <f> required to execute the turn 
■with the resultant load factor n is, of course, 


<f> — sin -1 — 
n 


( 7 ) 


By substituting the range r from equation (1) into 
equation (4), the centripetal acceleration at any azi- 

















290 


APPENDIX A 


r 



120 no 100 90 80 70 60 


Figure 1 . Scopodromic approach. 


50 

40 

30 

20 

10 

0 

10 

20 

30 

40 

50 







































INTERCEPTION VS STRAIGHT ESCAPE 


291 



Figure 2. Scopodromic approach curves with lines of constant load factor. 




























292 


APPENDIX A 


muth point a of a definite pursuit curve can be ex¬ 
pressed as 



sin 2 a 

tan ’f 


(4) 


At some place along the curve this acceleration be¬ 
comes a maximum. This critical place is defined by 
cos a. c = e/2, a remarkably simple results The criti¬ 
cal angle a c and the resultant load factor n max are 
shown on Figure 3. In a polar chart (Figure 4) the 



Figure 3. Load factor peak in horizontal scopodromic 
approach. 


critical azimuth is indicated by a straight line. It 
shows clearly how, for large speed ratios near 2:1, 
the banking is concentrated in the last phase of the 
combat, whereas for speeds nearly equal it occurs 
near a 60-degree azimuth. 

The magnitude of the horizontal load factor com¬ 
ponent peak depends also on the proximity of the 
approach, as expressed by the cross-path range r T , 
namely: 

T T / \ 1+ ‘/2 / vHJ 

= + * ( 1 _ t) (4a)d 

and the resultant load factor peak is then 

ttmax = \/l + W^nax (5a) 

d For proof see Section A.4.2. 


The critical range r c at which the maximum load 
factor occurs is, according to equation (1), 

r c - r r (sin a c ) (<-1) • (1 + cos« c ) _e (6a) d 


and the steepest banking angle 0 max IS 

(l+e/2) , v (l-«/2)_ 

•O-t) >"> 

All these values typically grow quickly with the 
product of the two speeds. Table 4 gives a c , n, ma x, 
Wmax, r c , and </> max for a v = 500-mph pursuer going 
after a bomber flying at speeds of V = v/e from 350 




Figure 4. Azimuth of greatest load, scopodromic. 


to 450 mph; and also for a 450-mph bomber with 
three different pursuit speeds. The proximity param¬ 
eter rr indicates the distance at which the pursuer 
passes abeam of the bomber. It is a parameter of the 
scale of the scopodromic curve and fully characterizes 
the path. 

Table 4, as well as Figure 2, gives a vivid picture 
of the load factor limitations of the pursuer's ap¬ 
proach. It is readily seen that no aiming (scopo¬ 
dromic) approach can be carried into firing range 
while still on a forward quadrant without entailing 
excessive load factors which would not only impair 
the aiming accuracy but also slow down the pursuer. 
I he critical azimuth a c , which is 563^ degrees for a 
speed ratio of 1/e = 0.9, is of some significance for 
the armament coverage of the bomber because from 
any greater angle the pursuer is hampered by the 
load factor's increasing as he approaches. Once the 
pursuer has negotiated this critical azimuth, the load 



















INTERCEPTION VS STRAIGHT ESCAPE 


293 


Table 4. Critical approach phase of interceptor in 
scopodromic pursuit. 


Bomber speed 

V (mph) 

Pursuit speed 

350 

385 

400 

450 

450 

450 

v (mph) 

500 

500 

500 

563 

500 

450 

Speed ratio \/e 
Critical azimuth 

0.7 

0.77 

0.8 

0.8 

0.9 

1.0 

angle 


44°21' 

49°32' 

51°19' 

51T9' 

56° 15' 

60° 

r T 








(yards) 








8,000 


0.863 

0.848 

0.85 

1.075 

0.867 

0.7313 


Wmax 

1.322 

1.311 

1.313 

1.468 

1.323 

1.239 


r c (yards) 

3184 

3856 

4104 

4104 

4820 

5333 


< t>m ax 

40°47' 

40° 18' 

40°22' 

47°4' 

40°56' 

36°10' 

4,000 

Wijmax 

1.725 

1.696 

1.70 

2.15 

1.735 

1.4625 


Wmax 

1.984 

1.969 

1.972 

2.37 

2.000 

1.77 


r c (yards) 

1592 

1928 

2052 

2052 

2410 

2666 


< t > max 

59°54' 

59°29' 

59°32' 

65°3' 

60°3' 

55°38' 

2,000 

Tljjmax 

3.45 

3.393 

3.40 

4.30 

3.47 

2.925 


Wmax 

3.59 

3.535 

3.545 

4.42 

3.62 

3.09 


r c (yards) 

796 

964 

1026 

1026 

1205 

1333 


< t > max 

73°50' 

73°35' 

73°37' 

76°54' 

73°55' 

71°8' 

1,000 

Wjjmax 

6.90 

6.785 

6.80 

8.60 

6.94 

5.85 


Wmax 

6.97 

6.850 

6.875 

8.66 

7.05 

6.65 


r c (yards) 

398 

482 

513 

513 

602 

666 


$ max 

81°45' 

81°37' 

81°38' 

83°22' 

81°48' 

80°18' 

factor 

eases up 

again 

, and he becomes 

more 

; dan- 


gerous. This need not mean that the tail guns of the 
bomber must cover a field extending 56 degrees off 
the tail axis, because the range at the critical point 
is still large; for our 500-mph vs 450-mph example, 
it is 850 yd for 5 g or 1,070 yd for 4gr, and the azimuth 
diminishes rapidly with further approach (to 30 de¬ 
grees when the range has shrunk to 512 yd for 5 g 
and 648 yd for 4# on the limit circle, or 655 yd for 5g 
and 825 yd for 4 g, respectively, along the critical 
approach curve). 

It is interesting to note that for a given pursuit 
speed, the maximum load factor suffered in scopo¬ 
dromic pursuit varies but little with the speed of the 
quarry, and is almost entirely determined by the dis¬ 
tance at which the approach arrives at any definite 
azimuth. As a matter of curiosity, it may be men¬ 
tioned that for a given pursuit speed, the least maxi¬ 
mum load factor is attained at e = 1.3, 1/e = 0.77, 
i.e., where the bomber’s speed is 30 per cent slower 
than the pursuit. e 


Time Elapsed 

The time it takes the pursuer to gain on the target 
from any initial azimuth angle is appreciable. Ac¬ 
cording to the simple scopodromic concept, at con¬ 
stant airspeeds, the elapsed time, counted backwards 


from the instant of catching up, is 



tan (<-1) ^ 


tan ( ‘ +1 >^ 
e t 1 _ 


( 8 )' 


The progress of the approach can thus be plotted 
on a time scale. This has been done in Figures 5, 6, 
and 7 for the example V = 450 and v = 500 mph. 
The plots refer to three initial positions (defined by 



Figure 5. Scopodromic approach against time; 
n max g = 3.000, rr = 2,455 yd, \f/ = 0. 



T IN SECONDS 

Figure 6. Scopodromic approach against time; 
Rmax0 = 4.000, r T = 1,793 yd, \f/ = 0. 


rr ) so chosen as to reach load factor peaks of 3, 4, 
and 5 g which occur 47}/£, 343^, and 273^ seconds, 
respectively, before catching up. The graphs show 
the gradual approach in range and azimuth toward 
the tail and the sudden surge of load factor and bank¬ 
ing angle. For instance, it appears that the intercep¬ 
tor has to be banked and releveled very accurately; 
the rolling speed has to attain peaks of the order of 
6, 83^, and 11 degrees per second. The maximum load 


See Section A.4.3 for proof. 


f See Section A.AA for derivation. 


















294 


APPENDIX A 


factors occur at ranges of 1,500, 1,080, and 840 yd, 
respectively. A factor halfway between 1 and the 
maximum is exceeded for 12.6, 9.6, and 7.9 seconds, 
respectively. Figure 8 is a synopsis, showing that the 
final approach phase proceeds practically independ¬ 
ently of the sharpness and proximity of the chase 
turn. 



Figure 7. Scopodromic approach against time; 
n m * x g = 5.00.9, r T = 1,419 yd, ^ = 0. 


Speed Loss 


However, it must be remembered that the idealized 
scopodromic approach is a theoretical simplification 
and the time history as shown here is but an approxi¬ 
mation of any real combat maneuver. For one thing, 
as has been pointed out, in reality the pursuer loses 
some of his theoretical speed advantage because, as 
he banks, the extra lift required to overcome cen¬ 
trifugal force entails induced drag, which slows him 
down. The order of magnitude of the retardation 
thus occasioned is 


D n 2 — 1 

g ' T ' x 4 + l 


(9)“ 


where D/L is the glide ratio at high speed, n is the 
load factor, and X = v/v (L/D)mmt is the ratio of the 
flight speed over his best glide speed. If, for example, 
D/L = 1 :14, X = 1.7, and an average n of 3, the 
deceleration would be 2 ft per sec per sec. This would 
consume all the speed advantage of the interceptor 
in 10 seconds, and half of it in the 5 seconds it would 
take him to creep in from 90 degrees to 56 degrees, 
the worst part of the approach. Allowance for this 
retardation would have to be made step by step if 
pursuit maneuvers of this character were to be 
studied in greater detail. As an additional refinement, 
the extra drag due to aileron deflection while chang¬ 
ing the bank angle may be taken into consideration. 


1 For proof see Section A.4.5. 


Vertical and Oblique Scopodromes 

Let us now allow the interceptor to attack from a 
different flight level so that he has to maneuver in 
space, and his flight path has curvature components 
both in horizontal and vertical projection, to be de¬ 
noted by the indices ?? and { in bomber’s flight co¬ 
ordinates. Theoretically, the mathematics of the 
scopodromic pursuit curve are the same for either 
component, as well as in the resultant slanting ap¬ 
proach plane defined by the pursued’s flight path and 
the line of sight between the two craft at some “ini¬ 
tial” condition. The main differences between the 
general case of oblique and horizontal approach are 
the load factor and the retardation suffered. 



Figure 8. Scopodromic approach, range against time. 


The total load factor n is now not merely y/n% + 1 , 
as for horizontal turns, but n — y/n% -f nj . With 
7 the angle of the slant-flight path plane against the 
horizon, this is determined by 


= sin 7 + y/l — si 


sin* a sin* 7 


+ (y cos y) (10) 


where c — FQ is the centripetal acceleration and 12 
the angular velocity of the interceptor’s axis in space 
during the oblique turn. 

The steeper the descent in the turn, the more se¬ 
vere is the extra load fact or because a given rate of 
turn produces only n — y /1 -f (c/g) 2 when c is hori¬ 
zontal and the path is horizontal, but n = 1 + c/g 
when c is vertical, and the path is horizontal. At a 
centrifugal force of 4/3 g the load increase gets twice 
as great in a vertical dive zoom (n = 7 / 3 ) as in a 
horizontal turn (n = 5/3). The higher the centrifu¬ 
gal force, the less difference the slope of the flight 
path makes. 

As to the retardation, the aerodynamical extra 
drag is also slightly greater for the descent turn in 
proportion to the effect of the increased resultant 













INTERCEPTION VS STRAIGHT ESCAPE 


295 


load factor n, but a gravity component — g sin y, 
helpful for negative y, readily offsets the induced 
drag. A descent of 4 degrees in the phase of the pre¬ 
viously adduced example* would suffice to balance 
completely the induced drag due to the turn. Steeper 
descents would actually help in boosting speed were 
it not for the bugaboo of the compressibility burble. 
In an attack from below, on the other hand, the 
gravity component becomes a serious retarding fac¬ 
tor, but the load factor diminishes slightly. 


Pitched Gun (Clinoscopodromes) 

In certain combat phases a gun installed at an ap¬ 
preciable angle to the flight path may have certain 
advantages (despite obvious ballistic and less ob¬ 
vious maneuvering disadvantages). If the gun were 
mounted at a skew angle, the phoronomy of the ap¬ 
proach would become mathematically rather com¬ 
plicated. The same is true if the gun were mounted 
at a fixed angle of pitch or yaw only, because as the 
interceptor banks, the displacement of the trajectory 
rotates around the flight-path tangent, and the latter 
degenerates into a bent skew spiral. The description 
of such a “helicodromic” path would require the solu¬ 
tion of two simultaneous differential equations be¬ 
tween the range and the pursuer’s position azimuth 
and latitude. Considerations of trajectory drop and 
angle-of-attack variation with load factor further 
complicate the mechanics of such an approach. If 
necessary, a step-by-step construction beginning with 
any given set of initial conditions can of course be 
executed, though it will at best be tedious and in¬ 
volved. « 

One particular case is amenable relatively easily to 
an explicit solution, namely, the attack in a vertical 
plane with a vertically elevated (or depressed) gun. 
In this “clinoscopic” case the approach does not 
require banking, and with a the apparent elevation 
of the pursuer as seen from the pursued and 0 the 
pitch angle of the interceptor’s gun with respect to 
its air path, the differential equation of the approach 
becomes 

dr _ e cog (3 — cos a (H) 

r € sin /3 + sin a 


For constant gun (and sight) inclination (3 this equa¬ 
tion is solved by 

€ COS d 


r = Tt 


a 

a — 1 + sin a 


1 + 


l - 



h For derivation see Section A.4.6. 


where a = 1 + e sin /3 and b = 1 — e sin 0. Here rr 
indicates that hypothetical range at which the pur¬ 
suer would have been vertically above (or below) the 
pursued (and upside down)) had he begun the ap¬ 
proach that far away. Some sample approaches are 
pictured in Figure 9 for a speed ratio 1/e = 0.9 and 
for gun inclination angles of 5, 10, 15, and 20 degrees 
up and —5 and —10 degrees down, load factor and 
azimuth being plotted against range. 

The load factor built up by virtue of the gradual 
“pull-out” is simply 


n = cos (a + P) — 


da 

dt 


which is 


n = cos (a + /3) + — • (sin a + e sin 0) (13) 

The loci of equal load factors are apple-shaped 
curves also shown in Figure 10. 

It is significant that, with the gun pitched, the 
pursuit terminates asymptotically in an approach 
from an angle a t slightly greater than the negative 
gun pitch angle /3, viz., sin a t = — e sin 0. During 
the end phase, the pursuer slowly creeps up from be¬ 
low (if gun is pitched up) or stalks from above (if gun 
is pitched down) with negligible path curvature and 
acceleration left at close firing range. 


A 3 2 Leading Pursuit (Ballodrome), 

Gun Parallel to Path 

It must be realized that all scopodromic pursuit 
maneuvers thus far studied, although automatically 
winding up in a stern chase, fail to make allowance 
for leading the target. In order to hit the target from 
such an approach while the azimuth is at all appre¬ 
ciable, the gun would have to be flexibly mounted in 
the interceptor with automatic or semiautomatic lead 
correction control. Obviously, the mechanical com¬ 
plication of such a system would be appreciable. It 
is much easier to apply the lead correction to the 
gunsight, either by automatic or semiautomatic con¬ 
trol, or even by “guess and experience.” This alter¬ 
native, however, complicates the mathematical treat¬ 
ment of the approach maneuver. It transforms the 
simple homing or scopodromic interception curve 
into one that, for the sake of descriptive identifica¬ 
tion, will be denoted as “ballodromic” (meaning “to 
drive so as to hit”). Its characteristic is that the in¬ 
terceptor does not head directly for the target but 
leads it by an ever diminishing lead angle in the plane 
of approach which is defined by the line of sight and 
in the target flight path (aside from the elevation 
correction for trajectory drop due to gravity). 









OL IN OEGREES 


296 


APPENDIX A 




2000 1500 1000 500 

r IN YARDS 


Figure 9. Clinoscopodromic approach. 

















130 

140 

150 

160 

170 

180 

170 

160 

150 

140 

130 


50 

40 

30 

20 

10 

0 

10 

20 

30 

40 

50 


INTERCEPTION VS STRAIGHT ESCAPE 



Figure 10. Clinoscopodromic approach with lines of constant load factor. 



































298 


APPENDIX A 


Range vs Azimuth 

If the gun elevation is controlled by a device at 
least compensating for airspeed (angle of attack) and 
trajectory drop so that the bullet trajectory chord is 
parallel to the straight flight path, then the lead cor¬ 
rection angle is determined by a “refraction” relation 

(14) 

sin a u v k € 


where u is the average bullet speed with respect to 
the muzzle or k is its ratio to the pursued’s airspeed. 
This can be directly derived from the bullet, inter¬ 
ceptor, and target speed triangle.* The polar differen¬ 
tial equation of the ballodromic curve is thus: 


dr 

r 


exp cot 8 — cot a 
1 — exp 


• da 


€\/esc 2 a — xp 2 — cot a 
1 — exp 


• da 


(i5 y 


The ballodromic equation is amenable to integration, 
so that the range can be expressed as a function of 
the azimuth by 


I [ cos 8 + xp cos a 

«*/2 

). 

LVcos 8 — xp cos a/ 

/ 

✓ \ -i 

/ COS 5 — COS a \ 

• A- 

NCOS 8 + COS a) 

sin a J 


Ranges with sin 8 = xp sin a from equation (14) 
have been computed for a representative value of 
speed ratio e and muzzle velocity ratio xp. The results 
are tabulated in Table 5 and shown in Figures 11, 12, 
and 13. 


Acceleration 

As to the banking angles and load factors, it is signifi¬ 
cant that in the ballodromic approach they are some¬ 
what less severe than in the scopodromic approach 
because the whole path is cut shorter. The entire 
maneuver takes several seconds less time. (The sco¬ 
podromic maneuver may be regarded as a special case 
of the ballodromic with infinite muzzle velocity, i.e., 

t = 0 .) 

An accurate evaluation of load factors n = sec <p 
in ballodromic approach (still with the gun assumed 
in the direction of the flight-path tangent) can be de- 


Table 5. Ballodromic approach. 

1/e = 0.9 V = 450 mph v = 500 mph r T = 1,419 yd 


Azi¬ 

muth 

a 

(de¬ 

grees) 

\f/ =0.225; u = 

= 2,200 fps 

xp = 0.183; it 

; = 2,875 fps 

Lead 

angle Bank 

8 Range 
(de- r (de¬ 

grees) (yards) grees) 

Load 

fac¬ 

tor 

n 

(g) 

Lead Load 

angle Bank fac- 

5 Range </> tor 

(de- r (de- n 

grees) (yards) grees) ( g ) 

150 

6.5 

32,300 

4 

1.00 

5.2 

20,870 

614 

1.01 

140 

8.3 

11,110 

15 

1.04 

6.8 

9,934 

1614 

1.04 

130 

9.9 

6,118 

28^ 

1.15 

8.1 

5,649 

31 

1.17 

120 

11.2 

3,806 

4334 

1.38 

9.2 

3,570 

4614 

1.46 

110 

12.2 

2,580 

56 

1.78 

9.9 

2,494 

58 

1.89 

100 

12.8 

1,865 

64 

2.29 

10.4 

1,840 

65 

2.38 

90 

13.0 

1,419 

6934 

2.83 

10.6 

1,419 

7014 

2.99 

80 

12.8 

1,124 

7234 

3.32 

10.4 

1,140 

7sy 2 

3.50 

70 

12.2 

921 

74 

3.67 

9.9 

941 

75 

3.87 

60 

11.2 

776 

75 

3.85 

9.2 

790 

76 

4.10 

55 

10.6 

719 

75 

3.85 

8.6 

743 

76 

4.06 

50 

9.9 

670 

75 

3.80 

8.1 

696 

7514 

4.00 

45 

9.2 

627 

74 

3.68 

7.4 

652 

75 

3.90 

40 

8.3 

589 

7334 

3.54 

6.8 

614 

7414 

3.73 

35 

7.4 

555 

7234 

3.32 

6.0 

581 

7314 

3.49 

30 

6.4 

525 

71 

3.05 

5.4 

550 

72 

3.22 

25 

5.4 

497 

6834 

2.73 

4.4 

522 

6914 

2.88 

20 

4.4 

470 

65 

2.38 

3.6 

496 

6614 

2.49 

15 

3.4 

433 

6034 

2.03 

2.7 

467 

6114 

2.08 

10 

2.2 

412 

51 

1.59 

1.8 

437 

5214 

1.64 

5 

1.2 

369 

3434 

1.21 

0.9 

394 

36 

1.23 

3 

0.7 

341 

24 

1.10 

0.6 

365 

25 

1.10 

1 

0.2 

290 

10 

1.02 

0.2 

313 

1014 

1.02 

H 

0.1 

245 

4 

1.00 

0 1 

198 

514 

1.01 


rived from the rate of change of the pursuer’s course, 
which is defined by the difference of the two azi¬ 
muths a — 5, viz. 

, v (da db\ 

«, = tan* = 7 ^-^j 

= — • (1 - • sin ail - if” —( 17 ) k 

<jr r \ cos 8 / v ' 

and hence the total load factor (if the approach takes 
place in a horizontal plane) is n = y/\ + n 2 . . 

The loci of equal load factors are ovals, slightly 
distorted from the scopodromic circles of equal load 
factors. These ovals and ballodromic approach curves, 
as they appear in a polar coordinate system flying 
with the bomber, are shown in Figure 11. The equa¬ 
tion for these ovals is 


vV (1 — exp) 
g\Zn* - 1 


• sin a 


cos 8 — \J/ cos a 
cos 8 


(18) 


1 Attack from below, however, cannot very well be accom¬ 
plished without loss of speed, whereas the assumption of no 
gain in speed in attack from above is not so far off because of 
the compressibility drag 

i For proof refer to Section A.4.7. 


Time Elapsed 

The time elapsed between passing any particular 
position ai and catching up has been determined for 


k For derivation and explanation see Section A.4.7. 


/ 































INTERCEPTION VS STRAIGHT ESCAPE 


299 


120 110 100 90 80 70 60 



50 

40 

30 

*20 

10 

0 

10 

20 

30 

40 

50 


Figure 11. Ballodromic approach with lines of constant load factor. 

























300 


APPENDIX A 


several sample conditions by graphical integration 
after having computed n v from equation (17) and 
according to equation (14) for suitable steps of azi¬ 
muth a. Figures 12 and 13 are time scale plots of 
ballodromic approach examples thus derived. These 
are interesting to compare with the corresponding 
scopodromic chart, Figure 7, which refers to the same 
“initial” position, namely, 1,419 yd abeam. While 
scopodromic pursuit from this initial position leads 
to a maximum load factor of 5, the ballodromic ex¬ 
amples reach 4.1 g and 3.85<7 for two different muzzle 
velocities, namely 2,875 and 2,200 fps, respectively. 
The critical azimuth angle at which the maximum 
load factor is reached is practically the same, 56 to 
57 degrees, independent of the aiming technique. 
The time elapsed from the abeam to the critical posi¬ 
tion is 3 seconds scopodromically and 4 seconds bal- 
lodromically; the time from there to theoretical col¬ 
lision is 273^ vs 24 seconds. 



40 30 20 10 o 

T IN SECONDS 


Figure 12. Ballodromic approach against time; 
= 3.85 g, tt = 1,419 yd, 0 = 0.225. 


Leading Pursuit (Ballodrome), 

Gun Fixed 

Thus far it has been assumed that the gun is auto¬ 
matically elevated to the tangent of the flight path 
of the interceptor. This assumption, however, is 
somewhat strained during the strongly banked phase 
of the ballodromic chase, where the high load factors 
induced by centrifugal force require larger angles of 
attack of the aircraft than in normal high-speed 
straight flight. With the gun fixed in the airplane, 
any change of angle of attack of the airplane is 


equivalent to an elevation or depression of the gun 
with respect to the flight-path tangent in straight 
flight as well as in turns. 


5g 5000 

4g 4000 

3g 3000 

S 

tt 

, * 

c z 

2 g 2000 

ig 1000 

0 0 

40 30 20 10 0 

T IN SECONDS 

Figure 13. Ballodromic approach against time; 
>W0 = 4.100, r T = 1,419 yd, 0 = 0.183. 



Influence of Angle of Attack 

The possible effect of this apparent gun-vs-path 
elevation upon the ballodromic approach phoronomy 
can be estimated. If the gun is fixed with the barrel 
axis parallel to the flight path in straight level flight 
at high altitude and high speed v and the excess angle 
of attack 6 of the aircraft varies essentially in propor¬ 
tion to the load factor, i.e., 6 = i(n — 1). The factor 
i is defined as the quotient of the lift coefficient of 
straight level flight to the slope of the lift-coefficient- 
vs-angle-of-attack curve of the interceptor; i is of the 
order of a degree or two. The maximum value of the 
excess angle of attack may be of the order of 6 = 5 
to 8 degrees, and this is about half of the maximum 
lead angle 5. 1 

In a horizontal turn the banking angle <f> is defined 
by n = sec <f>. The excess angle of attack has a hori¬ 
zontal component of r = Q sin 0 = i( tan 0 — sin 0), 
which is the angle to be deducted from the lead angle 
<5 to arrive at the flight path. The vertical component 
6 cos 0 = i( 1 — cos 0) may be disregarded; it would 
merely subtract itself from the trajectory drop cor¬ 
rection which, at a large banking angle, appears as a 
lateral or yaw displacement. Yet, an analytical cal¬ 
culation of the refinement, even under the further 


i 


Explained in Section A.4.8. 










INTERCEPTION VS STRAIGHT ESCAPE 


301 


simplifjdng assumptions of constant speed, leads to 
unmanageable expressions. 

Approach Between Scopo and Ballo 

As a first approximation it cannot be far off to ex¬ 
pect the resultant path and history of load factors to 
fall between the scopodromic and ballodromic curves 
corresponding to the same initial position, about half¬ 
way, in the phase of highest load factors, between 
a = 90 degrees and 30 degrees azimuth, and closer 
to ballodromic in the early (obtuse) and late (acute) 
phases. As a more accurate approximation a step-by- 
step correction method can be devised by determin¬ 
ing successive correction angles. (See Section A.4.8.) 

At any rate the maneuvers studied give a fair in¬ 
sight into the difficulties with which a high-speed 
chase at small-speed advantages is fraught. 


A 3 4 Intersection Passage Without Banking 

The approach maneuvers discussed in the preced¬ 
ing chapters are by no means easy to execute. They 
require an utterly precise banking technique, not 
only for the sake of preventing black-out but also for 
keeping the bead on the target. Thus, any type of 
approach requiring steep banking within the firing 
range may be impractical. Following is a study of 
what techniques of approach the interceptor could 
resort to in order to emancipate himself from the 
shackles of high load factors at firing range. Obvious¬ 
ly, the trick is for him to do most of his turning while 
he is still so far away that it can be done at moderate 
angular velocity and then to approach so that by the 
time he arrives within firing range he has sufficiently 
straightened out to minimize banking and load 
factors. 

Collision Course (Tomodrome) 

As a hypothetical maneuver, because of its aid in 
the treatment and presentation of the analysis, and 
not because of any particular tactical effectiveness, 
the study will first consider what may be called a 
“tomodromic” approach (meaning “driving so as to 
cut or intersect”), namely, leading to collision on in¬ 
tersecting straight-flight paths. This condition can be 
devolved from the ballodromic equation (11) by 
letting u = 0 or ^ = 1/e. This makes the denomina¬ 
tor zero, which indicates that the azimuth does not 
change when the proper lead 

sin 5 = — sin a (19) 

e 

is attained. 


Table 6 is a list of values of this proper lead angle 
for various azimuths and speed ratios. Figure 14 
shows them in a polar diagram. 


Table 6. 


a 



1/e 



a 


1/e 

(deg) 

0.5 

0.6 

0.7 

0.8 

0.9 

(deg) 

1.0 

1.1 

0 

0° 

0° 

0° 

0° 

0° 

180 

0° 

0° 

10 

4°59' 

5°59' 

6°59' 

7°59' 

8°59' 

170 

10° 

11°01' 

20 

9°51' 

11°50' 

13°5T 

15°52' 

17°56' 

160 

20° 

22°05' 

30 

14°29' 

17°27' 

20°29' 

23°35' 

26°45' 

150 

30° 

33°22' 

40 

18°45' 

22°05' 

26°45' 

30°56' 

35T9' 

140 

40° 

44°59' 

50 

22°50' 

27°23' 

32°25' 

37°48' 

43°38' 

130 

50° 

57°28' 

60 

25°40' 

31°20' 

37°18' 

43°52' 

51°16' 

120 

60° 

72°22' 

70 

28°01' 

34°20' 

41°09' 

48°41' 

57°34' 

110 

70° 


80 

29°30' 

36°14' 

43°35' 

52°00' 

62°23' 

100 

80° 


90 

30°00' 

36°52' 

44°26' 

53°08' 

64°09' 

90 




(A column for 1/e = 1.1 is included here to show what 
limited chance a slow interceptor has when his speed is 10/11 
that of his target; he has to attack from a forward sextant or 
else he would be left behind.) 

Tomodromic Approach Rate 

The actual rate of approach in the tomodromic 
maneuver is 

dr T ,| |/' /sina\ 2 sin 2 a\ 

~dt = F b- cos f 1 -(—) ~ — J 

= yV (20) 

Table 7 gives values for the bracket expression y 
which indicates the rate of approach as a fraction of 
the bomber’s get-away speed. Figure 15 is a hodo- 
graphic chart of it. 


Table 7. 


a 

(deg) 

1/e 

0.5 

0.6 

0.7 

0.8 

0.9 

1.0 

1.1 

0 

1.000 

0.667 

0.429 

0.250 

0.111 



10 

1.004 

0.671 

0.430 

0.251 

0.111 



20 

1.0155 

0.678 

0.435 

0.253 

0.111 



30 

1.037 

0.692 

0.443 

0.256 

0.112 



40 

1.0685 

0.710 

0.456 

0.262 

0.114 



50 

1.115 

0.745 

0.475 

0.274 

0.118 



60 

1.174 

0.791 

0.507 

0.290 

0.123 



70 

1.2565 

0.856 

0.553 

0.318 

0.133 



80 

1.364 

0.946 

0.624 

0.367 

0.151 



90 

1.500 

1.068 

0.729 

0.450 

0.211 



100 

1.660 

1.226 

0.976 

0.581 

0.325 

0.0602 


110 

1.860 

1.420 

1.069 

0.770 

0.499 

0.234 


120 

2.076 

1.645 

1.302 

1.010 

0.749 

0.500 

0.235 

130 

2.299 

1.887 

1.561 

1.351 

1.048 

0.826 

0.532 

140 

2.518 

2.130 

1.824 

1.576 

1.364 

1.173 

0.910 

150 

2.713 

2.344 

2.065 

1.844 

1.660 

1.500 

1.357 

160 

2.867 

2.518 

2.260 

2.060 

1.900 

1.764 

1.652 

170 

2.966 

2.635 

2.386 

2.201 

2.057 

1.940 

1.843 

180 

3.000 

2.667 

2.429 

2.250 

2.111 

2.000 

1.909 






















302 


APPENDIX A 


Slant Tomodrome 

Whether the plane of tomodromic intersection is 
horizontal or vertical, as in a glide approach from 
higher up or slant, the results are the same. Only the 
speed ratio is affected. The speed gain derived from 
a glide angle 7 at a given power is a matter of aero¬ 
dynamics near the compressibility burble limit and 
may have to be left to a step-by-step investigation of 
typical examples if of sufficient interest. 

A 3 5 Interception Passage (Brachydrome) 

The tomodromic approach would be suicidal if car¬ 
ried to the end, which is collision. It is theoretical 
also on the count that never during this approach is 
a fixed gun aimed at the proper lead angle; it would 
always fire ahead of the target. It is therefore necessary 
for the pilot, before he arrives at the desired firing 
range, to abandon the true tomodromic approach line 
and cut in behind the tail of his quarry. At close range 
he cannot afford to change over into the nearest ballo- 
dromic curve as this would entail an S-turn requiring 
extremely rapid aileron action, entailing high load 
factors and probably winding up in as much of a tail 
chase as a ballodromic approach from the beginning 
would have. To avoid this dilemma, all the pursuer 
has to do is to make an imperfect version of the 
straight slant approach, yet not to intersect but to 
miss the collision—in other words, slightly short. 
Such a maneuver will be denoted as “brachydromic” 
(“driving short”). This maneuver may attack the 
victim from vulnerable angles. However, the pur¬ 
suer’s firing rate must be very high, for he will have 
his guns aimed correctly onty during one instant of 
the passage and he may not be certain just at what 
range this will occur. 

The technique of this approach begins very much 
like the true intersection from initial azimuth a 0 , ex- 


120 HO too a *90 80 70 60 



310 330 o= <t 30 50 



Figure 15. Tomodromic approach rate. 


cept that the initial lead angle 8 0 is now chosen 
slightly less than the tomodromic lead angle of equa¬ 
tion (19). The choice of this new initial lead angle 5 0 
will be to a great extent a matter of training and ex¬ 
perience. Once it is attained, the interceptor pro¬ 
ceeds essentially straight and observes how the tar¬ 
get gradually creeps up toward the firing lead angle 
8* which is identified by the ballistic triangle, namely 
sin 8* = \p sin a*, while the azimuth off the target’s 
tail will have diminished slightly from a 0 to a*. The 
idea is to begin firing just before arriving at the 
proper range and lead and to continue firing until 
the target appears dead ahead. It must then have 
flown through a hail of bullets (provided proper al¬ 
lowance for trajectory drop was made). The pursuer 
need not swerve from his course and he is sure to fly 
through the wake of the victim, missing a collision. 

Brachydromic Approach 

The range equation governing the brachydromic 
passage is 

— = esin (0 -a Q ) + sin oc p ( . 

r 0 e sin (0 — a) -f sin a ^ 

where r 0 and a 0 are any given initial conditions and 
6 is the angle at which the two courses cross. If 
denotes the “shortness,” i.e., the range at which the 
enemy crosses dead ahead, then 

L = sing = 1 . _ sin ^_ 

o- e sin (0 — a) -f sin a e sin (a* — a) K 


















INTERCEPTION VS STRAIGHT ESCAPE 


303 


The rate at which the range changes is: 

^ = 7[cos a — e cos (a — 0)] 

and in bomber’s polar coordinates 

dr _ sin 0 [co s a — e cos (0 — a)] 
da [e sin (0 — a) + sin a] 2 

The azimuth a M at which the approach would have 
begun from r = oo is identified by: 

cot a x = cot 0 —— sin 0 ' (25) 

The closest approach ever reached occurs at an azi¬ 
muth of a 0 o — 90 degrees. 

Table 8 and Figures 16, 17, and 18 show how the 


(23) 

(24) 


range, expressed in terms of the dead-on “shortness” 
(which is not the closest passage distance), varies in 
the maneuver. 

In polar coordinates the p^ths appear as straight 
course lines intersecting the enemy course at the ini¬ 
tial azimuth angle a M while the pursuer is actually 
headed in the direction 0 as shown by the sample 
silhouettes. 


Brachydromic Firing Range and Lead 


The firing range at which the lead is correct is at¬ 
tained at a firing azimuth a* for which: 


tan a* 


_sin 0 

cos(0 — ^) 


(26) 


Table 8. Brachydromic interception approach at course intersection angle 0. 


a 


r/ 

<t for 1/e = 

0.5 




r/ 

<r for 1/e ■■ 

= 0.6 


(degrees) 

e = 30 c 

60° 

90° 

120° 

150° 

(degrees) 

30° 

60° 

90° 

120° 

150° 

0 

0.500 

0.500 

0.500 

0.500 

0.500 

0 

0.600 

0.600 

0.600 

0.600 

0.600 

10 

0.583 

0.507 

0.466 

0.422 

0.343 

10 

0.672 

0.597 

0.551 

0.498 

0.389 

20 

0.727 

0.531 

0.450 

0.375 

0.267 

20 

0.791 

0.595 

0.523 

0.437 

0.308 

30 

1.000 

0.577 

0.4485 

0.346 

0.224 

30 

1.00 

0.650 

0.515 

0.399 

0.257 

40 

1.685 

0.653 

0.460 

0.332 

0.198 

40 

1.415 

0.714 

0.521 

0.379 

0.237 

50 

6.10 

0.779 

0.488 

0.327 

0.183 

50 

2.55 

0.821 

0.532 

0.371 

0.208 

60 


1.000 

0.536 

0.333 

0.178 

60 

15.15 

1.000 

0.589 

0.375 

0.197 

70 


1.472 

0.615 

0.350 

0.172 

70 


1.33 

0.662 

0.391 

0.194 

80 


2.88 

0.751 

0.382 

0.174 

80 


2.085 

0.785 

0.413 

0.196 

90 



1.00 

0.433 

0.183 

90 


5.18 

1.000 

0.472 

0.205 

100 



1.565 

0.519 

0.199 

100 



1.438 

0.557 

0.221 

110 



3.91 

0.674 

0.221 

110 



2.71 

0.705 

0.244 

120 




1.000 

0.268 

120 



33.0 

1.00 

0.294 

130 




2.06 

0.345 

130 




1.86 

0.374 

140 





0.506 

140 




11.87 

0.536 

150 





1.000 

150 




. 

1.000 

160 






160 





9.62 

<*0O 

53.8° 

90.0° 

116.6° 

139.1° 

159.9° 

<*0O 

62.0° 

96.6° 

121.0° 

141.8° 

161.2° 



r/i 

t for 1/e = 

0.7 




r/a 

for 1/e — 

0.8 


(degrees) 

9 = 30° 

60° 

90° 

120° 

150° 

(degrees) 

30° 

60° 

90° 

120° 

150° 

0 

0.700 

0.700 

0.700 

0.700 

0.700 

0 

0.800 

0.800 

0.800 

0.800 

0.800 

10 

0.755 

0.684 

0.634 

0.572 

0.458 

10 

0.831 

0.765 

0.707 

0.639 

0.511 

20 

0.847 

0.688 

0.595 

0.495 

0.348 

20 

0.894 

0.756 

0.659 

0.543 

0.385 

30 

1.000 

0.721 

0.576 

0.450 

0.288 

30 

1.000 

0.769 

0.632 

0.472 

0.316 

40 

1.265 

0.766 

0.576 

0.423 

0.252 

40 

1.172 

0.808 

0.625 

0.456 

0.275 

50 

1.798 

0.854 

0.594 

0.411 

0.230 

50 

1.475 

0.881 

0.637 

0.443 

0.248 

60 

3.03 

1.00 

0.633 

0.412 

0.218 

60 

2.07 

1.000 

0.680 

0.442 

0.236 

70 

22.7 

1.25 

0.700 

0.426 

0.213 

70 

3.68 

1.197 

0.731 

0.453 

0.229 

80 


1.74 

0.810 

0.450 

0.215 

80 

18.5 

1.55 

0.833 

0.481 

0.231 

90 


3.02 

1.000 

0.505 

0.223 

90 


2.31 

1.000 

0.530 

0.240 

100 


12.92 

1.347 

0.588 

0.240 

100 


4.78 

1.303 

0.609 

0.257 

110 



2.41 

0.706 

0.269 

110 



1.95 

0.745 

0.286 

120 



6.56 

1.000 

0.316 

120 



4.15 

1.000 

0.335 

130 




1.67 

0.395 

130 




1.57 

0.418 

140 




5.59 

0.561 

140 




4.00 

0.582 

150 





1.000 

150 





1.00 

160 





5.32 

160 





4.00 

<*0O 

71.6° 

103.0° 

125.0° 

144.2° 

162.3° 

<*00 

82.5° 

108.5° 

128.7° 

146.3° 

163.3° 

































304 


APPENDIX A 


Table 8 ( Continued) 




r/i 

7 for 1/e = 

= 0.9 


(degrees) 

e = 30° 

60° 

90° 

120° 

150° 

0 

0.900 

0.900 

0.900 

0.900 

0.900 

10 

0.903 

0.837 

0.790 

0.709 

0.570 

20 

0.934 

0.821 

0.721 

0.601 

0.419 

30 

1.000 

0.821 

0.683 

0.535 

0.342 

40 

1.110 

0.846 

0.669 

0.496 

0.297 

50 

1.294 

0.903 

0.676 

0.476 

0.269 

60 

1.605 

1.000 

0.704 

0.471 

0.253 

70 

2.21 

1.16 

0.758 

0.481 

0.246 

80 

4.20 

1.43 

0.849 

0.507 

0.246 

90 

14.70 

1.94 

1.000 

0.554 

0.255 

100 


3.20 

1.262 

0.631 

0.272 

110 


11.7 

1.786 

0.764 

0.302 

120 



3.22 

1.000 

0.352 

130 



19.2 

1.51 

0.436 

140 




3.29 

0.598 

150 





1.000 

160 





3.35 

a* 

93.9° 

114.1° 

132.0° 

148.2° 

164.2° 

a 


r/o 

r for 1/e = 

1.0 


(degrees) 

30° 

60° 

90° 

120° 

150° 

0 

1.000 

1.000 

1.000 

1.000 

1.000 

10 

0.969 

0.922 

0.864 

0.777 

0.612 

20 

0.969 

0.879 

0.780 

0.653 

0.451 

30 

1.000 

0.866 

0.732 

0.577 

0.366 

40 

1.065 

0.879 

0.710 

0.532 

0.315 

50 

1.180 

0.922 

0.710 

0.508 

0.285 

60 

1.367 

1.000 

0.732 

0.500 

0.268 

70 

1.668 

1.13 

0.780 

0.508 

0.259 

80 

2.28 

1.35 

0.864 

0.532 

0.259 

90 

3.73 

1.73 

1.000 

0.577 

0.268 

100 

11.1 

2.53 

1.233 

0.653 

0.285 

110 


4.98 

1.67 

0.777 

0.315 

120 



2.73 

1.000 

0.366 

130 



8.14 

1.462 

0.451 

140 




2.88 

0.612 

150 





1.000 

160 





2.97 

a* 

105.0° 

120.0° 

135.0° 

150.0° 

165.0° 

a 


r/o 

for 1/e = 

1.1 


(degrees) 

e = 30° 

60° 

90° 

120° 

150° 

0 

1.10 

1.10 

1.10 

1.10 

1.10 

10 

1.03 

0.995 

0.936 

0.850 

0.66 

20 

1.00 

0.935 

0.843 

0.701 

0.582 

30 

1.00 

0.908 

0.777 

0.615 

0.388 

40 

1.03 

0.908 

0.746 

0.563 

0.336 

50 

1.10 

0.937 

0.741 

0.537 

0.301 

60 

1.21 

1.000 

0.757 

0.524 

0.282 

70 

1.40 

1.108 

0.799 

0.530 

0.272 

80 

1.73 

1.285 

0.875 

0.552 

0.273 

90 

2.34 

1.585 

1.000 

0.596 

0.280 

100 

3.57 

2.16 

1.21 

0.669 

0.297 

110 

11.1 

3.55 

1.59 

0.788 

0.328 

120 


10.96 

2.42 

1.000 

0.379 

130 



5.49 

1.423 

0.464 

140 




2.61 

0.624 

150 




18.8 

1.00 

160 





2.72 

Of 00 

115.1° 

124.7° 

137.7° 

151.6° 

165.7° 


and the lead angle at this instant is 8* = a* — 6 for 
which: 


sin 8* = \p sin a* or tan 8* = 


\J/ sin 6 


1 — \J/ cos 6 


(27a) 


or 

tan 8* = 


tan 8 


1 H— • cos 8 

v 


[cos 5± j/^iy — sin 2 8 


(27b) 


where u/v is the ratio of the muzzle velocity to the 
interceptor’s own airspeed. 

When the pursuit pilot does not know and cannot 
estimate the course of the quarry, he can,theoretically, 
so maneuver that he finds at what lead or crab angle 
8 he can fly straight without the target changing its 
apparent angular position during the approach. He 
knows his own bullet-to-flight speed ratio u/v. This 
would suffice to solve the ballistic triangle, except 
that the inaccurately known enemy-to-own speed ra¬ 
tio 1/e = V/v enters in equation (27b) as a slight 
correction. 111 The resulting ballistic lead angle re¬ 
quired to hit the target is, however, ambiguous. The 
larger one corresponds to the case of encounter, the 
smaller one to that of overtaking. The values of 
proper lead angles for various tomodromic approach 
crab angles 8 t and a series of bullet-to-pursuit speed 
ratios u/v = 2.7, 3.0, and 3.3 times the pursuit speed 
are given in Table 9 and plotted in Figure 19 and a 
cross plot for u/v = 3 in Figure 20. 

Table 10 and Figures 21 to 25 give a synopsis of 
values of proper lead angles and ranges in brachy- 
dromic interception for a series of course intersection 
angles 6 in steps of 30 degrees and for various speed 
ratios e and \p. 


Example 


To quote an example, take the case of \f/ = 0.225, 
which may correspond to a v = 500 mph interceptor 
(1/e = 0.9) and u/v — 3 (effective average muzzle 
velocity u = 2,200 fps). If this interceptor pilot 
wanted to so intercept a V = 450 mph bomber as to 
hit at, say, r* = 400 yd, he would have but a few 
seconds for the several phases of such an attack, and 
less than a second of effective fire. Table 11 shows 
for various course angles 6 how much time 


r* sin_5* _ r* sin a* 
V sin 0 u + v sin 0 


(28) 


A torpedo-directing technique can be developed from this 
method of tomodromic approach and brachydromic inter¬ 
ception. 


> 































INTERCEPTION VS STRAIGHT ESCAPE 


305 


06 


0.5 


0.6 


330 


340 


350 


10 


20 


30 



Figure 16. Brachydromic passage, 1/e = 0.6. 



























306 


APPENDIX A 


06 0.7 I 06 0.8 

330 340 350 0 10 20 30 



Figure 17. Brachydromic passage, 1/e = 0.8. 
































INTERCEPTION VS STRAIGHT ESCAPE 


307 



Figure 18. Brachydromic passage, 1/e = 1.0. 




























308 


APPENDIX A 


Table 9. Proper brachydromic lead angle 8* after 
tomodromic approach at crab angle d. 



u/v = 

2.7 

u/v = 

3.0 

u/v = 

3.3 

8 

Si* 

St* 

Si* 

St* 

Si* 

5,* 




For 1/e 

= 0.7 



10° 

3°49' 

i°r 

3°34' 

0°56' 

3°21' 

0°51' 

20° 

7°18' 

2°9' 

6°49' 

1°57' 

6°22' 

1°47 # 

30° 

9°56' 

3°33' 

9°1T 

3°13' 

8°37' 

2°58' 

40° 

10°50' 

5°50 / 

10°01' 

5°20' 

9°18' 

4°54' 

44°26' 

8°51' 

8°51' 

8°05' 

8°05' 

7°28' 

7°28' 




For 1/e 

= 0.8 



10° 

3°57' 

0°42' 

3°42' 

0°38' 

3°24' 

0°35' 

20° 

7°36' 

1°28' 

7°07' 

1°20' 

6°40' 

1°13' 

30° 

10°36' 

2°23 / 

9°52' 

2°10 / 

9°14' 

1°59' 

40° 

12T8' 

3°39' 

11°25' 

3° 18' 

10°37' 

3°02' 

50° 

11°35' 

6°06' 

10°39' 

5°33' 

9°51' 

5°05' 

53°08' 

8°55' 

8°55' 

8°13' 

8T3' 

7°24' 

7°24' 




For 1/e 

= 0.9 



10° 

4°5' 

0°22' 

3°50' 

0°20' 

3°37' 

0°18' 

20° 

7°54' 

0°45' 

7°24' 

o°4r 

6°57' 

0 o 37' 

30° 

11°09' 

1°13' 

10°26' 

1°06' 


1°00' 

40° 

13°23' 

1° 47 / 

12°28' 

1°37' 

11°35' 

1°28' 

50° 

14°02' 

2°4T 

12°57' 

2°25' 

12°01' 

2°12' 

60° 

11°52' 

4°27' 

11°01' 

4°02' 

9°57' 

3°40' 

64T0' 

7°45' 

7°45' 

7°or 

7°or 

6°24' 

6°24' 




For 1/e 

= 1.0 



10° 

4°13' 

0° 

3°58' 

0° 

3°44' 

0° 

20° 

8°12' 

0° 

7°41' 

0° 

7°14' 

0° 

30° 

11°39' 

0° 

10°53' 

0° 

10°13' 

0° 

40° 

14°15' 

0° 

13T7' 

0° 

12°24' 

0° 

50° 

15°35' 

0° 

14°25' 

0° 

13°23' 

0° 

60° 

15°10' 

0° 

13°52' 

0° 

12°51' 

0° 

70° 

12°21' 

0° 

11T4' 

0° 

10°19' 

0° 

80° 

7°12' 

0° 

6°30' 

0° 

5°54' 

0° 

90° 

0° 

0° 

0° 

0° 

0° 

0° 


elapses between the arrival at proper lead angle 8* 
and zero lead (dead-on) and how far the bullets of 
any one machine gun firing at the rate of, say, 
6 = 20 rounds per second would be spaced on the 
moving target, along the target’s longitudinal axis 

* = y • * (29) 

and laterally with respect to the sighting line 

«y = y • t sin a* (30) 

Bullet Density 

It is immediately apparent that for any intercep¬ 
tion passage at appreciable crossing angle 6 only one 
or two bullets per gun will have a chance of hitting 
a vital part of the airplane. The fact remains that the 
slanting or brachydromic attack can be pressed home 


Table 10. Values of r*/<r for + = 0.175, 0.200, 0.225, 
and 0.250, and for 6 = 30°, 60°, 90°, 120°, and 150°. 
l/€ = 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, and 1.10. 


e = 

0° 

30° 

60° 

90° 

120° 

150° 

180° 


^ = 0.175 

a* 

0 

35°54' 

69°27' 

99°55' 

127°58' 

154°2r 

0 

5* 

0 

5°54' 

9°27' 

9°55' 

7°58' 

4°21' 

0 

1/e = 0.5 

1.27 

1.315 

1.425 

1.56 

1.69 

1.78 

1.81 

0.6 

1.175 

1.205 

1.31 

1.43 

1.56 

1.64 

1.675 

0.7 

1.10 

1.14 

1.235 

1.35 

1.47 

1.55 

1.57 

0.8 

1.055 

1.09 

1.185 

1.30 

1.41 

1.48 

1.505 

0.9 

1.025 

1.06 

1.15 

1.26 

1.37 

1.43 

1.46 

1.0 

1.00 

1.035 

1.12 

1.23 

1.34 

1.40 

1.425 

1.1 

0.98 

1.015 

1.10 

1.21 

1.31 

1.37 

1.40 


<£ = 0.200 

a* 

0 

36°55' 

70°54' 

101T8' 

128°56' 

154°51' 

0 

’ 8* 

0 

6°55' 

10°54' 

iris' 

8°56' 

4°51' 

0 

1/e = 0.5 

1.33 

1.38 

1.53 

1.70 

1.85 

1.95 

2.0 

0.6 

1.20 

1.24 

1.37 

1.53 

1.66 

1.76 

1.80 

0.7 

1.12 

1.17 

1.29 

1.43 

1.55 

1.64 

1.68 

0.8 

1.07 

1.11 

1.22 

1.36 

1.48 

1.56 

1.61 

0.9 

1.03 

1.07 

1.18 

1.31 

1.43 

1.51 

1.55 

1.0 

1.00 

1.04 

1.15 

1.27 

1.39 

1.47 

1.50 

1.1 

0.98 

1.02 

1.12 

1.24 

1.36 

1.44 

1.47 


^ = 0.225 

a* 

0 

37°58' 

72°22' 

102°41' 

129°56' 

155°25' 

0 

8* 

0 

7°58' 

12°22' 

12°41' 

9°56' 

5°25' 

0 

1/e = 0.5 

1.41 

1.48 

1.65 

1.87 

2.05 

2.19 

2.23 

0.6 

1.24 

1.30 

1.45 

1.64 

1.80 

1.94 

1.96 

0.7 

1.14 

1.20 

1.34 

1.51 

1.66 

1.78 

1.81 

0.8 

1.08 

1.13 

1.26 

1.43 

1.57 

1.68 

1.70 

0.9 

1.03 

1.08 

1.21 

1.37 

1.50 

1.61 

1.64 

1.0 

1.00 

1.05 

1.17 

1.32 

1.45 

1.55 

1.58 

1.1 

0.97 

1.02 

1.14 

1.29 

1.42 

1.51 

1.54 


* = 0.250 

a* 

0 

39°05' 

73°53' 

104°02' 

130°54' 

155°53' 

0 

5* 

0 

9°05' 

13°53' 

14°02' 

10°54' 

5°53' 

0 

1/e = 0.5 

1.50 

1.59 

1.76 

2.08 

2.30 

2.47 

2.50 

0.6 

1.28 

1.36 

1.52 

1.77 

1.47 

2.11 

2.14 

0.7 

1.17 

1.24 

1.38 

1.61 

1.78 

1.92 

1.95 

0.8 

1.09 

1.15 

1.30 

1.50 

1.67 

1.79 

1.82 

0.9 

1.04 

1.10 

1.24 

1.43 

1.58 

1.70 

1.73 

1.0 

1.00 

1.06 

1.19 

1.37 

1.53 

1.64 

1.67 

1.1 

0.97 

1.03 

1.16 

1.33 

1.48 

1.59 

1.62 


if the quarry pursues a straight course. The success 
of this type of interception is, however, very prob¬ 
lematical because it depends on (1) rapid fire power 
to assure that enough bullets will hit the target, 
(2) prompt alignment on an interception path that 
will bring him within the desired range, and (3) early 
detection and identification of the enemy and his 
flight direction. 

The first condition is necessitated by the fact that 
the bullet stream sweeps or saws through the target. 
Ihis may be an advantage inasmuch as it mitigates 






















































INTERCEPTION VS STRAIGHT ESCAPE 


309 


Table 11. Firing opportunity for r* = 400 yd, b = 20 
rounds per sec, * = 0.225, v = 500 mph, V = 450 mph, 
u = 2,200 fps, s = 25 ft. 


e 

Time to dead-on 
ti (sec) 

Lateral spacing 
s v (ft) 

30° 

0.52 

15 H 

60° 

0.45 

23 K 

90° 

0.40 

24 

120° 

0.35 

183^ 

150° 

0.29 

9 


any inadvertent horizontal aiming error, but it may 
also be a weakness if the bullets are spaced too far 
apart, especially in a near broadside attack where the 
scatter is greatest. The interceptor must, of course, 
be equipped with several guns adapted to be fired 
syncopatingly. 

As to the second problem of prompt alignment, the 
maneuver is more difficult than an early phase of a 
scopodromic turn. The trick is to straighten out into 
a near tomodromic intersection approach and to cor¬ 


rect the interceptor's course quickly so that the crab 
angle (0) appears to remain nearly constant until the 
firing range is approached. Then, instead of swerving 
toward the wake of the target, the interceptor need 
merely throttle slightly to let the target creep up 
close to the proper lead angle 8*, which can be de¬ 
rived from the previously observed crab angle 0 ac¬ 
cording to Table 9 and Figure 19, if only the speed 
ratio e can be estimated with sufficient accuracy. As 
soon as the target has arrived at the proper lead an¬ 
gle 8*, or slightly before, the interceptor may open 
the throttle wide again and fire. 

It will be noted that the final maneuver is free from 
turning and involves no increased load factors while 
in or near the firing range, but the actual combat time 
is much shorter than in the ballodromic tail chase. 
The question remains whether there is enough time 
to determine the constant crab angle 0 and the cor¬ 
responding firing lead angle 5*. The computation of 
the latter can be automatized by means of a cam de¬ 
vice set to the proper value of the speed ratio e as 












310 


APPENDIX A 



0 20 40 60 80 100 

S IN DEGREES 

Figure 20. Brachydromic lead, for u/v — 3.0. 


soon as it is estimated. The pilot would merely adjust 
a diopter toward the target during the tomodromic 
approach phase with constant crab angle. This would 
automatically cam the proper correction into his re¬ 
flector gunsight, ready for use when he is ready to 
“cut short.” 

However, it is possible that the maneuver can be 
learned and practiced even without such a device by 
allowing for the lead to diminish from the tomo¬ 
dromic value which is sure to be too large, to zero 
(dead-ahead) which is sure to be too small. The prob¬ 
ability of a hit, however, remains small at best. 

The third problem, prompt detection and identifi¬ 
cation of the quarry, is aggravated for the interceptor 
by the shortness of the time available (1 to 2 min¬ 
utes) to obtain and digest any information received 
from the ground, and the time (but a few seconds) 
to maneuver into any other than a tail chase position. 

Slant Brachydrome 

The brachydromic interception can be executed in 
a sloping path as well as in a horizontal plane. The 


descent yields a speed advantage which, though 
small, is constant and easy to compute. 

A36 Outlook 

The protagonist of minimum defense on the fast 
bomber may concede that interception from blunt 
or even obtuse angles is possible but he may justly 
argue that such attacks will be rare, because of the 
early decision required to prevent the encounter from 
developing into a stern chase, and that at best they 
will be weak or ineffectual because of the brief time 
of combat fire. Should the first attack miscarry, then 
a second one by the same attacker would inevitably 
wind up on the tail. The more blunt or obtuse the 
first attack, the greater the chance of the bomber to 
run away because of the speed drop of the pursuer 
in his turn to resume the chase. In fact, the latter 
may risk losing his quarry altogether. For example, 
if the paths crossed at 90 degrees and if the pursuer 
made an immediate quarter-turn at 4# with his speed 
dropping down to the same (450 mph) as the bomb¬ 
er’s, he would drop about 1,300 yd behind. On the 




INTERCEPTION VS STRAIGHT ESCAPE 


311 



Figure 21. Brachydromic firing range, \f/ = 0.175. 



















312 


APPENDIX A 


330 340 350 0 ^ 10 20 30 



Figure 22. Brachydromic firing range, ^ = 0.200. 

















INTERCEPTION VS STRAIGHT ESCAPE 


313 


330 340 350 0 ^ tO 20 30 



Figure 23. Brachydromic firing range, = 0.225. 


















314 


APPENDIX A 



40 


50 


60 


70 


80 

90 

100 


110 


120 


130 


140 


210 200 190 180 170 160 150 

Figure 24. Brachydromic firing range, \f/ = 0.250. 



















MATHEMATICAL DERIVATIONS 


315 


other hand, if an acute tail attack, either brachy- 
dromic or ballodromic, is unsuccessful, all the pur¬ 
suer has to do is to zoom up, drop behind, and resume 
the tail chase. The tail chase remains the normal 
combat phase of the fixed gun interceptor going after 
a high-speed, high-altitude bomber flying straight. 

What now if the interceptor is equipped with some 
flexible guns with which he can attack obliquely from 
some angle? Such an installation may be considered 
as belonging to the category of a “new weapon” but 
it cannot well be entirely ignored. (In fact, it is not 
entirely new; in World War I, the Nieuport fighter 
was thus equipped.) The old argument that slant 
fire is inaccurate because it requires compound lead 
correction is not wholly conclusive. If the pursuer 
succeeds in creeping up into some position on top or 
below or off to the side or at some skew corner of his 
victim and then flies parallel to it at the same speed, 
he can leisurely pour lead into it, keeping an uncor¬ 
rected bead on it since the lead corrections for the 
two parallel motions would cancel each other. The 
approach phase was treated in Section A.3.1 under 
Pitched Gun (Clinoscopodrome), but once the inter¬ 
ceptor approaches the terminal position angle he 
may throttle back and stalk his quarry. 

The only obvious arguments against the effective¬ 
ness of this stalking technique are: 

1. The approach into such a dangerous position 
would probably come from the tail through a clino- 
scopodromic or ballodromic pursuit during which the 
attacker would have been exposed to the tail defense 
of the bomber. The latter’s job would be to prevent 
the pursuer from ever catching up with it. Maneuver¬ 
ing into a skew or top or bottom stalking position on 
a tomodromic interception approaching from an un¬ 
defended blunt angle would require a great deal of 
skill and practice. 

2. In firing at blunt angles the projectiles are likely 
to tumble, especially at the extremely high airspeeds 
contemplated here. 

3. To aim obliquely while flying is difficult for any 
pilot, who, after all, has to keep one eye on where he 
is going, especially at high speed. The interceptor 
may carry a separate gunner for the oblique gun, but 
even so the pilot has to watch the quarry closely 
enough to fly parallel to him. It is a task similar to 
formation flying at high speed. It is reasonable to 
assume that this technique is therefore limited to 
moderate cone angles off the tail of the pursued, per¬ 
haps up to 45 degrees. Attack from the upper quad¬ 
rant would be particularly difficult because of vision 
limitations; attack from underneath the tail would be 
least difficult. 

In speculating about “new weapons” that may be 
developed to bolster the defense against faster and 
higher-flying bombers—aside from stepping up the 



Figure 25. Lead angle for brachydromic interception. 


firing speed of the guns—the thought of immunizing 
the interceptor personnel against higher acceleration 
stresses, rather than avoiding them, is worth follow¬ 
ing through. By seating the pilot in a crouched posi¬ 
tion with the body bent forward and the face for¬ 
ward, the limit of resistance to impairment of vision 
may be extended 2 g, from Qg to 8 g for 2 seconds or 
from 5g to 7g for somewhat longer periods. On the 
other hand, a reclining position, with the head back 
and as low as possible, compatible with a clear view 
through the gunsight, and with the feet raised to 
pedal extensions and legs suitably supported, may be 
even more effective in reducing the total head of 
blood and yet be more comfortable and helpful in 
attaining the steadiness required in aiming, though 
special gunsights may have to be devised for this 
technique. Liquid pressure suits may eventually be 
perfected to raise the acceleration tolerance. How¬ 
ever, similar concessions to the protection of the per¬ 
sonnel might also be considered for the bomb er to 
raise its maneuvering limitations, insofar as they may 
be governed by physiological rather than structural 
or stall limits. 

a.4 MATHEMATICAL DERIVATIONS 

A 41 Scopodromic Pursuit Curve 

in the Pursued’s Polar Coordinates 

Analysis 

Radial speed component: 

dr TT 

—r = V cos a — v. 

dt 







316 


APPENDIX A 



v = true airspeed of pursuer 
€ = v/V = speed ratio 
a = azimuth of pursuer off pursued’s tail 
r = range 

rr = range in cross-path position 
0 = refers to any initial position 

Figure 26. Diagram of scopodromic pursuit. 

Azimuthal speed component: 

da _ _ V sin a 
dt ~ r 

Quotient of above: 

dY _ _ V cos a — v 
da V sin a 

Rearrange and substitute e = v/V: 

— = A -cot a)da. 

r \ sin a / 


Transform: 

d In r — d In tan — d In sin a. 
Integrate: 

In r = In tan* ^-In sin a + C. 


Antilog: 

tan *f 

r = c • —:-. 

sin a 


Same for initial position: 



r 0 == c • . -. 

sin a 0 


Quotient : 


tan 


, «o 

tan T 


sin ocq 
sin a 


and for r 0 = r T , «o = 90 degrees: 

a 


tan* 


r = rr- 


sm a 


Q.E.D. 


Peak Load Factor 
in a Scopodromic Curve 


Analysis 


d(n v g) _ 


da 


= 0 


for n„g = 
and a — a c . 


vV sin 2 a 


r T , , OL 

tan* 


vV_ 

r T 


sima 


( 1 + COS a ) 

\1 — COS a) 


<1 2 


For a quotient, the derivative vanishes when the 
product of its numerator by the derivative of the 
denominator equals the product of the denominator 
by the derivative of the numerator; that is, 


sin 2 a c d ^tan* tan* °^- d(sin 2 a) 


Thus, 


sup a c 


da 


tan* -1 -=r- 


da 


2 cos 2 


2 a 

- = tan* • 2 sin a c cos a c 


Everything cancels except 


Thus, 

and 

Hence, 


~2 — (‘os a c 


cos- a c = ~r 

4 f 2 
sin 2 a c = 1-— 


Q.E.D. 


sin J 


/l + cos a c V /2 _ / 

i A 

r + n 

\1 — COS a c J \ 

1 - t) 

i.-j 


72 


■ (‘ + i) o-i) ( 1 + t) 
-('+*).•(>-*) 


«/2 


1 —«/ 2 


(>-t) 

Q.E.D. 






















MATHEMATICAL DERIVATIONS 


317 


Equation (6a) can be transformed to read 
r c = r T (l - y) 0- + *) _< 


Least Load Factor Peak 
in Scopodromic Pursuit 


The maximum centripetal acceleration is 

l+e/2 / v l-t/2 


vV 

rrg 


(-i) -(-f) 


and for v = constant, V = v/e, the above becomes a 
minimum for 


d 

ro+ir - ( 

i- 

I 

* |<N 

1 

rH 

de ’ 

e 

* V l+e/2 


Let v 1 + T/ =a 

and 0 -t) 

and 

Then: 

dA , dB D 

~j — = A and -j— = B 
de de 

(BA' + AB')e - AB = 0 

Here, however 

A' — \A In | 

(■+t) 

and 

B' = \B In 

(>-t) 


so that 


minimum and not a maximum is readily seen, when 
the second derivative is inspected, namely, 


f k 1 - t ) ^ k 1 + t ) 





2 


i 



This is indeed positive (which is the condition for the 
presence of a minimum) for all values of e < 2. (Of 
course, for larger speed ratio no load factor peak is 
reached during the approach phase, as the load factor 
would keep on increasing until the pursuer has caught 
up with his victim; no a c can have a cosine greater 
than 1.) 


A 4,4 Time of Scopodromic Approach 

From Section A.4.1, the reciprocal of the azimuthal 
speed component is 


dt r 

da V sin a 


and from equation (1), a second form of 


tt 

sin a 



— cos a 

+ COS a, 


e/2 


When combined, they result in 


dt 

da 


r T 

V sin 2 a 



*[lBA In (l + -j) ~ h^B • In (l - y)]= AB, 
where AB cancels, and by combining the logs, 


The order of magnitude of its solution is readily 
found by expansion and with 



and the actual value by a little trial and error. It 
furnishes c/2 = cos a c = 0.65, a c = 49 degrees, and 
c = 1.3; 1/e = 0.77. That the extreme value is a 


Now, since sin 2 a = 1 — cos 2 a= (1 — cosa)(l + cosa) 

d^ _ _ rr (1 — cos a) t,2 ~ 1 
da V (1 + cos a) fl2+1 


This is integrated by the substitution 


which makes 

and 

and 



x 


1 


cos a = 


2x 2 

1 + x 2 


1 


+ cos a = 
da = 


2 

1 + x 2 
2 dx 
1 + x 2 


With these substitutions the integral becomes 


t = 


r T 

2V 



( x f ~ 2 + x e )dx 

















318 


APPENDIX A 


which is 


t = - 


r T 

2V 


x e_i x e+l 1 

E - 1 + 6 + lj 


r T 




2FL € - 1 1 € + 1 


Q.E.D. 


Hence, after substitution and cancellation of tt e/R, 

C D _C*h - Ch 
C* D 2 C*h 


Now, in order to carry the same weight, 

r _c* L 

Thus, 

Cd — C*d • "b 0~ 4 ) = Co(l + <t>~ A ) 


A - 4 5 Deceleration Due to Inertia Load 


Nomenclature 


V(L/D) m „ is the airspeed of best glide angle. 

<f> = v/v( L / D ) m „ is the ratio of the actual speed to 
that of the best glide angle. 
e/R = effective aspect ratio. 


_ C 2 l 

Ci = is the induced drag coefficient. 
Co is the parasite drag coefficient. 


Analysis 

At any speed the drag coefficient is made up of 

C 2 l 


Cd = C o + 


ir e/R 


At best glide angle (indicated by *) it is: 

C* L 


C* D = Co + 


ire/R 


but here n 


Co = C*i = 


C* 2 l 

■k e/R 


n Proof for the fact that, at the best glide angle, induced and 
parasite drag are equal: 


where e/R is the effective aspect ratio. 

"»• t-%* a 

The derivative of this with respect to Cl, must vanish: 

1 


,Cd 

1£l „ c. . 

dC L C* L ve/R 


for least 
namely, at 
Hence, 


Cd 

Cl 

Co_ 1 
C* L ive/R 

Co = — n. 

Co ice/R C * 


Now in the presence of a load factor n, the induced 
drag increases from Ci — C 0 /<£ 4 to n 2 C 0 /</> 4 . The dif¬ 
ference in drag coefficient then becomes A Cd = AC, 
= C 0 (n 2 — 1 )/<£ 4 and its ratio to the steady-flight 
drag Cd becomes 

A C D _ n 2 — 1 
Cd 1 + 0 4 

With S = wing area and q = velocity head, the 
weight carried at high speed is 


Mg = C L Sq 


so the deceleration of the same mass M due to excess 
drag is 


-i = AC D •§ = 


A Cd 

~cV 


g 


A C D (Cp\ 

C d \C l J ♦ ' g 

( D\ n 2 — 1 

? \T/V0M r T 


Figure 27 illustrates a clinoscopodromic approach. 




f Light 

Figure 27. Diagram of clinoscopodromic approach. 


Clinoscopodromic Approach 

Analysis 

Radial speed component: 

dr Tr 

= V cos a — v cos /3 

Azimuthal velocity: 

da _ F sin a — v sin (3 
dt r 













MATHEMATICAL DERIVATIONS 


319 


Quotient of above, with v = eV 

dr _ 6 cos (3 — cos a 
da e sin j8 + sin a 


whence equation (11), which is (with dr/r => d In r) 

J , € COS 0 — COS <2 

a In r = —:— : - • da 

e sin 0 + sin a 

To integrate, since € and /I are constant, split 

>/; 


In r = e cos 0 , 


e sin jS + sin a 


-U 


COS a da 


e sin jS + sin a 
The last of these integrals is simply 
In (e sin 0 + sin a) 


+ C 


The first one (according to Manual of Mathematics 
and Mechanics by Clements & Wilson, integral 205 b) 
is 


* 2 

y/l — e 2 sin 2 j8 


tanh -1 • 


y/l — e 2 sin 2 ft 
. 1 + € sin /3 



Ballodromic Approach 


Figure 28 illustrates a ballodromic approach. 



v = pursuer’s true airspeed 
V = pursued’s true airspeed 
€ = V/V 
# = V/(u - f- v ) 
r = range 
a = azimuth 
5 = lead angle 

Figure 28. Diagram of ballodromic approach. 

Analysis 


With a = 1 + e sin /3, and b = 1 — e sin /3, this is ab¬ 
breviated to: 

^• tanh - l [|/|- tan (f“T)] 

However, for the sake of getting rid of the other log¬ 
arithms it is preferable to express the tanh -1 by a 
logarithm also, according to tanh -1 z = In (1 + z) 
/(I — z), where the [ ] expression above plays the 
role of z. 

Thus, 

In r = 6 C —^ • In ] — - — In (e sin (3 + sin a) + C 
y/ab l - z K 

and for a = 90 degrees, where tan (a/2 — x/4) = 0 
so that z = 0, 

In r T = —In (e sin /3 + 1) + C 
The difference of the two is free from C ; its antilog is: 

(e cos 0) 

r _ e sin (3 + 1 T 1 + z ~l %/ab 

r T e sin /? + sin a Ll — zj 

which turns into equation (12) by substituting for 
1 + e sin (3 and for z. 


According to the law of sines in the interception 
speed triangle ABC: 

sin 8 _ V _ 
sin a u + v ~ * 

Speed of range change AE — BF, 
d r ir 

-jt — V cos a — v cos o 
at 

Rate of azimuth change (FD — EC)/AB , 
da _ —V sin a -f v sin 5 
dt r 


Their quotient is 
dr _ 
da 


or 



V cos a — V COS 8 

----- • y 

V sin a — v sin 8 

cos a — e cos 8 

— -:—- • r 

sin a — e sin 8 


cos a — € cos 8 
sin a — e sin 8 


• da 


Substitute for 8 


dr _ e's /1 — </> 2 sin 2 a — cos a 
r sin a(l — e*) 


• da 


ey/esc 2 


cot a 


1 — ei/' 


• da 


























320 


APPENDIX A 


Hence 


ln r = ~ 1 ' da 
- I cot a da = 

The second integral is simply / 2 = In sin a + C 2 . 
The first one is solved by the substitution 0 
xp COS a xp COS a _ EA 

EB 


\/ \ — xp 2 sin 2 a 


n ^ /v COS b 

and since sin 5 = xp sin a 

the differential is 

da = — 


V 1 - V- 2 


(1 - z 2 )V^ 2 - 2 2 

so that the first integral, h, is transformed into 

dz 


Ii = (V 


-»/< 


{V - **)(1 - *) 

This is split into a product of linear denominators by 
+ xP-Z 


+ -°- + JL 

' 1 “I" 2 ' 1 


1 


Z (xp 2 — Z 2 )(l — Z 2 ) 

and solved by successively eliminating all but one 
term after another by successively allowing z = —xp, 
xp, — 1,1, which furnishes A = B = %xp(l — xp 2 ) and 
C = D = y&(xp 2 — 1). With these constants the solu¬ 
tion becomes 

-Hfef -m\ 

When this is introduced into the original problem 

[ / \*^ 2 / \ *^/ 2 

ln (*r;) (ffi)- lnsina 

For a = 90 degrees, where 0 = 0 the range is r = r T . 
Thus, 

Resubstitution for 2 now furnishes equation (16). 

Q.E.D. 


0 In view of the complete equivalence of the respective 
azimuths a and 5 of one ship against the other, it is logical to 
seek a solution of the part of the integral which involves the 
lead angle by referring the geometry not to the coordinate 
system of either ship but to the only system that is equivalent 
to both, namely, their instantaneous line of sight. The pro¬ 
jections on it of target speed and bullet speed bear the ratio 
z = t cos tt/cos 5 and by virtue of the “refraction” law, 
sin 5 = ^ sin a, governing the relation between the two 
angles, this is also the ratio of the two azimuth tangents, viz., 
z = tan 5/tan a. 


The calculation being somewhat cumbersome, it 
may be of interest to note that approximations will 
serve to speed up the computation where second- 
order terms can be neglected. Inasmuch as xp is of the 
order of 20 to 25 per cent, its second-order terms are 
practically negligible, especially in view of other er¬ 
rors or inaccuracies of assumptions, notably the dis¬ 
regard of the variation of the angle of attack of the 
aircraft with varying load factor. Therefore, a first- 
order expansion approximation is useful to gain an 
insight into the influence of ballodromic leading, for 
any special case; that is to say, 

dr ^ e esc a — cot a , 


which can be directly integrated to give 

l/(l-«*) r —* 1 / [ 2(1—] 

(1 — COS a) 4-1 

- r l(l + COS a) <+I J 

which significantly turns into the scopodromic for¬ 
mulas (2) when xp is neglected. 

To visualize the significance of the exponent, it is 
well to remember that it is half of 1/(1 — exp) 
= 1 — v/u, in which v/u is the ratio of the pursuer’s 
own airspeed to his own gun’s average effective 
muzzle velocity. 

For the last phase of the stern chase, when the 
higher-order terms of a also can be neglected, the ap¬ 
proximate range equation boils down to 


L sin a J 


r 

r 0 



(«-l)/(!-«*) 


/ a\ <i-v/i>) 
\OCoJ 


which again agrees with the scopodromic case, equa¬ 
tion (3), as soon as xp is disregarded. 

For the calculation of load factors, take da/dt from 
the third equation of this section. 

For db/dt observe that: 


and since 


db _ db da 
dt da dt 


cos b 


sin b 
db 


da 


xp sin a 
xp cos a 


hence 


db 

da 


= t 


cos a 
cos b 


Thus, 


da db 
dt dt 


V 

r 


(1 


. /. . COSa\ 

exp) sin all — xp -- 1 

V cos b) 


To determine the time elapsed, the reciprocal of 
the apparent azimuthal angular velocity, dt/da 
= r/V sin a(l — exp), is plotted against a, and the 
graph is integrated by planimeter, which yields the 

































MATHEMATICAL DERIVATIONS 


321 


time elapsed. This was done for e = 10/9 and two 
values of \f/ = 0.183 and 0.225. 

A 4 8 Influence of Angle of Attack 

The influence of angle of attack is shown in Fig¬ 
ure 29. 



Figure 29. Diagram of angle of attack. (For a. = 90 
degrees, sin 5 = thus, for ^ = 0.225, 8 = 13 degrees.) 


In computing the progress of the approach step by 
step, the following procedure was adopted. 

From an initial position chosen far enough away 
so that the influence of the angle of attack is negli¬ 
gible, determine a first approximation of n from the 
ballodromic load factor from equation (17). Thence, 
compute 6 — i{n — 1). For a practical example the 
aerodynamic coefficient i may be assumed 1 de¬ 
grees. Now lay off the line of flight at a course 
a — 8 + r (r = 0 sin </>) and progress at the rate v 
for a step of sec. The new position furnishes new 
values of a and r and Aa/At from which a new bank¬ 
ing angle is computed according to 


V 2 e r . i 

tan <f> = - e sin (<$ — r) — sin a I 

gr 


whereupon the process is repeated. 

During the subsequent scopodromic approaches, 
the azimuth a decreases, thus shifting the peak of the 
frequency curve more toward the acute angles. Each 
point of the frequency curve is shifted from the ini¬ 
tial azimuth a 0 to a new azimuth a x corresponding 
to the same scopodrome, which is defined by a certain 
value of tt satisfying both a 0 and r 0 . The value of the 
ordinate, however, undergoes a change from the 
original value of p a0 to p al because the same number 
of attacks now cover a different range da, such that 


Poda 0 — pidai; hence pi = p 0 da 0 /dai. The corre¬ 
sponding “breadth’’ da 0 and da i are bordered by two 
scopodromes of slightly different parameters r T - 
Therefore the variation da 0 /dai must be expressed by 
the ratio of dao/df(r T ) and dai/df(r T ). 

The parameter rr is defined according to equation 
(1) by any set of corresponding values of r and a, 
namely, 

r sin a 

r T = - 

tan 'f 

Its derivative with respect to the azimuth is 


drr COS a — e COS a — e 

= r -= r T -•—— 

da a Sin a 

tan 2" 

Hence, 

da sin a 


d In r T cos a — e 

The ratio of the two corresponding derivatives 
thus becomes: 

da o _ e — cos sin «o 
dai e — cos «o sin a\ 

and with this 

e — cos a\ sin ao 

Pi = Po • - • —- 

e — cos ao sin ai 

In other words 

sin a 

p • --= constant 

e — cos a 


In order to determine the variation of p at the 
theoretical limits of a = 0 and 180 degrees, note that 
pi/'po approaches a 0 /«i for a -*■ 0 and (180 — a 0 )/ 
(180 — ai) for a 180. Now the first-order terms 
of the series expansion of r = rr tan e 3^a /sin a are 
r T oL e ~ 1 /2 e and r T 2 e /(180 — a) <+1 , respectively. The 
ratios of the corresponding limit angles thus become 
tied up with the ratio of the initial and subsequent 
range by 



for a -> 0 and 180 degrees, respectively. In view of 
this, the limit probability ratios become: 

1/C«—1) . vl/(«+l) 

and pt = 


01 = Po (f) 


respectively. 

Note that for speed ratio e = 10/9 the increase of 
the limit probability at tail azimuth a\ = 0 is af¬ 
flicted with an exponent 9 and quickly assumes astro¬ 
nomical proportions. The probability of receiving 














322 


APPENDIX A 


fire from any finite azimuthal sector, though less 
spectacular, also crowds most markedly toward 
a = 0. 


A5 COMBAT MANEUVERING TECHNIQUES 

The point has been made that the chances of an 
interceptor’s destroying a straight-flying, ultra-high¬ 
speed, high-altitude bomber from any quarter but an 
acute cone around the tail are remote and that it 
would therefore appear justified and economical to 
provide the bomber with, defense for the tail cone 
only. 

What now if the enemy became wise to such lim¬ 
ited defense and if he developed the technique and 
practice of slant (brachydromic) attack to the point 
of a menace? The bomber might then be forced to 
change his tactics, abandon the straight-course flight 
plan and dodge the interceptor. 

It is at once obvious that a drastic change of alti¬ 
tude would avail the bomber nothing, as the bomber 
and interceptor would suffer similar performance 
changes. Any loss of altitude would only aggravate 
the danger of later interception by other fighters. The 
following study will therefore first be directed to the 
effect of veering maneuvers in a horizontal plane. We 
shall consider three distinct types of maneuvers: 
luring the interceptor into the tail cone (see Section 
A.5.1), foiling the interceptor’s attack by spoiling his 
aim or lead (see Section A.5.2) and heading toward 
the interceptor to deprive him of maneuvering time 
(see Section A.5.3). 

A 51 Veering Away to Force Tail Combat 

The most obvious maneuver to thwart an attack 
from an undefended angle is to veer away from the 
interceptor. This automatically brings the intercep¬ 
tor into the tail cone where strong defense is assumed 
available. Any attempt on the part of the interceptor 
to avoid the tail-cone defense zone would of neces¬ 
sity spoil his aim, increase his range, and practically 
ruin his chances. 

Aerodynamic Ballistic Advantage 

If the interceptor’s forward guns are of the same 
fire power, caliber, and muzzle velocity as the tail 
defense guns of the bomber, then the bomber is at a 
great advantage over the interceptor, for the follow¬ 
ing reasons: (1) The bomber’s tail guns being flexible 
(though within limits), the bomber pilot need not 
aim the entire airplane whereas the fighter pilot must 
do just that; at high speed this is much more difficult 
than at conventional speed because of the high ac¬ 


celerations or inertia forces accompanying every con¬ 
trol movement. (2) The bomber’s rearward fire is 
more accurate and the projectile has greater impact 
energy left because of the lesser trajectory drop and 
lesser air resistance as the projectile speed against air 
is u — V for the rearward fire versus u 4- v for for¬ 
ward fire; the ratio of these bullet speeds is almost 
1:2 so that the air resistance of the attacker’s bullets 
is several times that of the bomber’s. This is im¬ 
portant because at the larger ranges at which combat 
will probably have to begin at the higher flight speed, 
trajectory drop is quite pronounced. As an example, 
the deflection of a .50-in. bullet is tabulated in Table 
12 as computed by extrapolation from Aberdeen 
Proving Grounds Ballistic Research Laboratory Re¬ 
port No. 117 for various ranges and for ilq = 2,700 fps 
initial muzzle velocity and for firing backwards from 
a bomber flying 450 mph and forward from an inter¬ 
ceptor flying at 500 mph. 

Table 12. .50-in. bullet deflection and drop. 

wo = 2,700 fps initial muzzle velocity 
V = 450 mph bomber firing rearward 
v = 500 mph fighter firing forward 
h = 40,000 ft altitude 


Range 

(yards) 

Deflection 

Trajectory drop* 

Energy ratio 

Fighter 

forward 

(mils) 

Bomber 

rearward 

(mils) 

Fighter 

forward 

(feet) 

Bomber 

rearward 

(feet) 

Bomber : 
Fighter 

600 

4.6 

4.1 

8.35 

7.4 

1.124 

800 

6.4 

5.7 

15.3 

13.6 

1.159 

1,000 

8.3 

7.3 

24.8 

22.0 

1.202 

1,200 

10.3 

9.0 

37.2 

32.5 

1.251 

1,400 

12.4 

10.95 

52.0 

46.0 

1.318 

1,600 

14.7 

12.9 

70.4 

61.8 

1.370 


‘Computed as a first approximation by extrapolation from Aberdeen 
Proving Ground Ballistic Laboratory Report No. 117, assuming resistance 
to vary approximately proportional to air density and square of bullet flight 
speed. 

Figure 30 is a graph of these values versus range. 
At lower altitudes the difference is even much more 
pronounced because of the greater air density. 

Similar tables can be constructed for 20- and 37- 
mm caliber if trajectory data are made available. 

Disparity of Arms 

If the interceptor carries guns of larger caliber and 
/or much greater muzzle velocity than the bomber, 
then this may offset the interceptor’s handicap. The 
bomber pilot will then have to resort to other tactics 
to shake off the attacker. He can still veer away from 
him as long as he is out of the attacker’s range in 
order to gain time, especially if the speed differential 
is small. However, it will be essential for the bomber’s 













COMBAT MANEUVERING TECHNIQUES 


323 


commander to know the armament and performance 
characteristics of the attacker. The bomber com¬ 
mander may even decide to enter a mock dogfight 
and emerge from it into tail combat at a short enough 
range to bring his own tail defense armament into 
most effective action. 

A52 Dogfight 

If the bomber wishes to avoid tail combat, he can 
accept a dogfight before the attacker has approached 
to within his firing range. 

Attempts to express the phoronomy of such a dog¬ 
fight in analytical terms indicate that the results are 
too complicated and cumbersome to evaluate load 
factors and lead angles. Even in the simple idealized 
case of the bomber flying in a steady circle and the 
pursuer following in a scopodromic spiral, the equa¬ 
tions describing the pursuer’s path are rather un¬ 
manageable. However, some insight into the effect of 
various maneuvers can be gained by graphical con¬ 
struction of the pursuer’s path. 

Acceleration Handicap of Pursuer 

It is immediately apparent that if the pursuer con¬ 
tinues at full power scopodromically, ballodromically, 
or somewhere in between after the bomber has turned 
toward his side in front of the pursuer, then the 
pursuer’s path tightens up and reaches a much higher 
load factor than the steadily turning bomber. 

If the pursuer is unable to stand more than a cer¬ 
tain load factor in combat, say 4 or 5, then he has to 
relinquish his quarry and let it pass. 

The success of an escape turn on the part of the 
bomber depends a lot on the ratio of the turning 
radius to the range at which the turn is begun. If the 
radius is large compared to the initial range, then the 
chase might develop into an advantage for the pur¬ 
suer who would essentially trail the bomber, just 
slightly cutting short to properly lead the target. 
Thus the pursuer would catch up eventually and his 
load factor would be but slightly higher than that 
suffered by the bomber—in fact, little more than in 
proportion to the square of his speed advantage. 

If, however, the turning radius is commensurable 
to the initial range, i.e., if the bomber does not let 
the attacker approach closer than a couple of thou¬ 
sand yards, then the attack can possibly be out- 
maneuvered by turning toward the attacker. If this 
maneuver is judiciously executed, it can be made to 
lead to a close-range encounter passage in which, it is 
true, the attacker has an exceedingly brief chance of a 
burst, but under exceedingly unfavorable aiming con¬ 
ditions. Immediately afterwards, however, the bomb¬ 
er’s tail defense has a chance of hitting the pursuer 


under much better aiming conditions, safe from re¬ 
turn fire. The next phase following the encounter of¬ 
fers the bomber a gain of range before the pursuer 
can turn around and resume the chase. The bomber 
may so maneuver that the pursuer blacks out in the 
chase or, if the bomber does not want to pass a given 
certain load factor, he loses his pursuer before he ar- 



Figure 30. Impact energy ratio and bullet flight time. 

rives at firing range. This sort of maneuver effectively 
shakes the pursuer off while the bomber passes ahead 
of him. If the pursuer attempts an S turn after pas¬ 
sage, then the bomber will straighten out its course 
when in opposite position and put so much distance 
between himself and the pursuer that the engagement 
is broken off. 

Tracking in Circular Flight 

The pursuer may, of course, prefer to follow the 
pursued in his track rather than cut short across the 
turn and avoid the higher load factor in the later 
phases. He would then simply creep up behind him 
though it would take a little longer. However, this is 
easier said than done. For one thing, the flight path 
does not usually remain visible, and when it does 
leave a condensation track, it is preferable not to fly 
through it. Secondly, when tracking along a curved 
path, the aim is very far off, unless a very large gun 
elevation is available. Obviously, tracking 1,000 yd 




324 


APPENDIX A 


behind on a 1,000-yd radius places the target 30 de¬ 
grees off the pursuer’s flight-path tangent, to which 
a lead correction still has to be added. 

Initiative of Escape 

Once the bomber has taken the initiative and 
started to turn in the direction to force the higher 
load factor upon the pursuer, the sense of turn in the 
dogfight must not be changed because the first one 
to make an S turn suffers a tactical disadvantage, ex¬ 
cept when the bomber decides to accept tail combat 
in a flight direction favorable to him with regard to 
the sun, clouds, or reinforcements. 

A 5 3 Head-on Parry 

If the interceptor is detected while still in a for¬ 
ward quadrant, then the bomber may choose to head 
directly for the attacker. Even though the latter may 
not be flying at top speed, the range will now dimin¬ 
ish so fast (say at 400 to 450 yd per sec) that only 
one to two seconds are available for combat. 

Snag Dodge 

The bomber allows the attacker to come on to 
within about 2,000 yd, or almost within long-range 
firing distance, and then veers out of his way to spoil 
his aim. An effective escape maneuver now consists 
in an S snag so close to the enemy that he has no 
time to turn after the bomber and cannot turn sharp¬ 
ly enough to lead properly. Then the bomber gets 
away before the pursuer can complete a 180-degree 
turn. The chances of a destructive hit in such a de¬ 
layed veering maneuver from a head-on encounter 
are very slim indeed. 

The interceptor may choose to attack ballodromic- 
ally during the encounter but he will get only a few 
rounds in and these under very unfavorable lead con¬ 
ditions, practically at “cross paths,” which means 
about 20-degree lead. (At 600 yd, this is 6 to 8 bom¬ 
ber’s lengths.) During or immediately after the frus¬ 
trated close-range passage, the bomber reverses his 
banking angle and returns to his original course head¬ 
ing. While the pursuer completes his more than 180- 
degree turn, he loses about 3,000 yd in distance be¬ 
fore he can resume the tail chase which will bring him 
within a few degrees of the bomber’s tail by the time 
(almost 2 minutes) he may again arrive within firing 
range, unless he has lost his quarry in the melee. 

Head-on Passage 

On the other hand, the interceptor may choose a 
second alternative, namely, that of foregoing all 


chance to hit or fire at the bomber during the first en¬ 
counter, and passing the bomber’s course ahead of 
him merely in order to lose less distance for a subse¬ 
quent tail attack. This, however, the bomber can foil 
by starting to trail the interceptor. The trick here 
will be to avoid being dragged away from the original 
objective or lured into the fire of other fighters. 

A 5 4 Vertical Escape Maneuvers 

The question may be raised if vertical escape ma¬ 
neuvers or their combination with horizontal ones 
have merits for the bomber. 

Not in Tail Chase 

In the tail chase the answer apparently is no. Most 
likely, the pursuer would suffer the lesser load factor 
and he would be in a favorable position whenever the 
bomber levels off. 

Looping 

Only if the bomber were to start looping just be¬ 
fore the pursuer has arrived at his maximum firing 
range would the attack be foiled. However, the pur¬ 
suer could probably follow suit and be right on the 
bomber’s tail after the loop is completed or when the 
bomber rolls out at the top of the loop. To continue 
looping would theoretically be a possible defense tac¬ 
tic but it is hardly practical and would only serve to 
let other fighters catch up and join the chase. 

Zooming Up Behind 

The pursuer, however, may benefit from vertical 
maneuvers. For instance, if in catching up fast he 
fails to bring his quarry down, he may instead of 
veering out of the way, zoom up behind him to kill 
speed only to utilize the potential energy thus gained 
in diving after the quarry in a subsequent attack. 
However, the only effect of such a maneuver as com¬ 
pared with merely throttling would seem to be the 
interruption of fire which may be desirable while 
changing ammunition drums or clearing jams. Other¬ 
wise, between a fixed gun pursuer and a limitedly 
flexible tail gun of the bomber, the advantage would 
usually be on the side of the latter. 

Descent for Lower-Altitude Bombing 

The bomber may be tempted or forced to descend 
to lower altitudes to fulfill his mission when poor vis¬ 
ibility obscures the target at great height. During the 
descent itself he does not necessarily enhance his risk 
because in the descent he picks up extra speed. The 



MULTIPLE INTERCEPTION 


325 


proximity of the compressibility limits his gain, just 
as it does to the interceptor who may be trailing him 
down. The interceptor may perhaps gain a little more 
if it has thinner wings but this advantage may be 
balanced by the fact that the interceptor is probably 
already designed to have its best performance at a 
lesser lift coefficient than the bomber so that the 
same speed-up would increase its drag more. At the 
higher descent speeds, load factors in any curved pur¬ 
suit increase with the square of the true speed so that 
the hazards of angular fire are still further reduced 
until the advantage is absorbed by the greater air 
density. 

The time the bomber has to stay at lower altitudes 
can be very brief. The descent can be made very flat 
and may take 50 miles, certainly enough to correct 
for a reasonable navigation error. After delivery of 
the bombs, the bomber’s climbing speed is increased 
and he may still have excess kinetic energy to re¬ 
trieve several thousand feet before being slowed 
down to the steady climb rate. During the remainder 
of the climb to the stratosphere he would have little 
to fear from those interceptors that might be taking 
off to go after him. Only in case other interceptors 
have been hovering above and dive after him would 
he be at a speed disadvantage. However, he could 
have known of their presence if he was equipped with 
sufficiently long-range detecting apparatus and could 
have decided to stay up until he might have shaken 
off or outlasted the interceptors or chosen an alter¬ 
native target for which he would not have to descend. 

A55 Recapitulation 

To summarize the results of the maneuvering 
study thus far, it may be said: 

The best plan undoubtedly is to provide the bomb¬ 
er with sufficiently powerful tail armament so that 
it can accept a tail chase on better than equal terms 
with whatever interceptor is fast enough to catch up 
with him before running out of fuel. Better than 
equal terms does not necessarily mean more power¬ 
ful weapons because of the aerodynamical bullet 
speed advantage for the rear fire, as has been ex¬ 
plained in Section A.5.1 under Aerodynamic Ballis¬ 
tic Advantage. 

At a speed differential of only 50 mph (or 25 mph) 
it takes the tail chaser fully 5 minutes (or 10 min¬ 
utes) to creep up from 7,000 yds into firing range. 

Any attack from a forward quarter can be turned 
into an almost head-on encounter and foiled short of 
firing range. 

The bomber can so maneuver that any attack will 
eventually wind up as a tail chase and that the at¬ 
tacker will drop behind whenever he tries any other 
trick. The bomber cannot be sure of shaking a tail 


chaser off permanently, but he can probably turn any 
angular attack into a tail combat whenever he wants 
to. 

The bomber can accept a dogfight and may out¬ 
last the interceptor. 

A 5 6 Maneuvers with Slant Gun 

The most effective countermeasure to improve the 
interceptor’s chances in a dogfight near the load fac¬ 
tor limits would appear to be its provision with some 
elevated guns. This will permit the interceptor to fire 
across a chord of the dogfight circle without running 
into excessive load factors—without, in fact, having 
to cut too much across the victim’s path. In such a 
sharp turn, both ships are steeply banked so that the 
gun elevation becomes essentially an azimuth correc¬ 
tion. The slight lateral correction, consisting of the 
gun elevation times the sine of the banking angle, can 
be compensated by the pursuer circling slightly lower 
than the pursued and by trajectory drop. Visibility 
for aiming such a gun elevated at, say, 15 degrees can 
be easily provided for the pilot. This arrangement 
will also come in very handy in a stalking straight 
attack from below the tail without dogfighting. How¬ 
ever, such an elevated gun requires means for proper 
lead correction for the own-speed vector, and it does 
not by any means assure superiority of the pursuer 
against a bomber equipped with tail guns covering 
similar cone angles and similar lead correction de¬ 
vices. 

The consequences of the presence of a slant gun on 
the interceptor have already been mentioned. If a 
high-speed fighter version of the high-speed bomber 
prototype were eventually developed, it might to 
good advantage also be equipped with a slant gun of 
moderate caliber or one that can be elevated in 
turns. 

A6 MULTIPLE INTERCEPTION 

The question now arises whether concerted pur¬ 
suit by several interceptors having fixed guns may 
become unduly dangerous to a lone bomber. 

A 61 Multiple Brachydromic Approach 

If an interceptor squadron flying in close formation 
tries to attack the fast bomber in brachydromic ap¬ 
proach from some odd angle, the individual intercep¬ 
tors would be forced to peel off as the rear ones would 
have to head more ahead of the bomber. Therefore, 
if all the interceptors were to participate in the fire 
their formation would become loosened up. As for the 
bomber’s defense, the situation hardly differs from 
the brachydromic attack by a single interceptor be- 



326 


APPENDIX A 


cause the individual fighters come into range one by 
one and all from the same quarter, almost from the 
same angle. 

A 6 2 Multiple Tail Chase 

If several interceptors were to come up from the 
rear and tried to place themselves simultaneously 
around the tail—say one to the right, another to the 
left, one above and another below—and if they were 
now to press into firing range together or in such 
rapid syncopation that the bomber’s defense gunnery 
could not cope with all of them, then the bomber 
might veer slightly to the side of the nearest attacker 
just before he comes into firing range. This maneuver 
brings all the attackers into the same quarter, and 
they would lose a lot of time if they rearranged them¬ 
selves. 

a.s.s Multiple Ballodromic Approach 

In ballodromic approach a squadron of intercep¬ 
tors would be more handicapped by high load fac¬ 
tors, aggravated by the requirements of maintaining 
formation, than a single interceptor. Otherwise the 
situation for the bomber’s defense winds up in a tail 
chase or can be turned into a tail chase prematurely 
whenever the bomber veers to run away. 

A 6 4 Multiple Clinoballodrome 

A situation annoying to the bomber can possibly 
arise from simultaneous attack by two pursuers ap¬ 
proaching ballodromically, but one in or above and 
the other below the bomber’s level. Of course, they 
would quickly get into each other’s way if both were 
equipped with fixed-level guns only. However, if at 
least the lower one is equipped with an elevated gun, 
they can attack simultaneously without interference. 
The bomber might then try to turn in order to ham¬ 
per the two closely flying attackers by high load fac¬ 
tors. Such a tactical turn may be especially indicated 
where the two attackers make it a habit to fly in a 
staggered formation, the one above slightly ahead of 
the lower one, leaving the burden of avoiding inter¬ 
ference to the lower one who has the better visibility. 
In the escape turn, the upper forward one becomes 
the inner one and gets into the firing line of the rear 
lower one, which has to take the outer line, and there 
is not much the latter can do about it but cease firing. 

A 6 5 Simultaneous Interception 

Now what if several interceptors were to make con¬ 
certed attacks from entirely different sides and angles 
timed to arrive simultaneously? Such tactics are 


highly improbable. It would seem almost impossible 
to spot the bomber so accurately, to disseminate the 
information to all concerned, and to work out and 
transmit a timed multiple interception plan to the 
several fighter units, all of which were moving at a 
rapid pace while these preparations were taking 
place. Then they would have to approach from sta¬ 
tions several miles apart in space. The least error in 
their calculations or any deliberate course change on 
the part of the bomber would completely upset the 
schedule. It seems certain that most of the intercep¬ 
tors, insofar as they do not miss the engagement al¬ 
together, would wind up in a tail chase. The small 
speed differential would make it difficult and tedious 
for them to sneak up at once. 

A 6 6 Multiple Frontal Attack 

Multiple frontal attack is dismissed as impractical 
because of lack of maneuvering room and time. 

A 7 MASS RAID TACTICS 

Navigating a fleet of bombers together toward a 
common objective or toward different objectives may 
offer each bomber some measure of protection. 

A 71 Decoy Action 

To what extent such protection may be secured by 
decoy action, drawing the interceptors to a feint and 
away from the real raid, is a matter of strategy which 
is considered outside the scope of the present study, 
except insofar as it may have a bearing on the arma¬ 
ment requirements for the bomber. If some of the 
decoy bombers were equipped with extra armament 
instead of full bomb loads, a bluff might be perpe¬ 
trated which might discourage the enemy from at¬ 
tacking the real bombers from undefended angles or 
ranges. Such extra armament should then be so de¬ 
signed that it is not easily distinguishable and that 
the standard and specially armed aircraft cannot be 
told apart in the air. 

Lateral Attack 

In a mass formation raid by a fleet of fast high 
bombers, dodging of interception would be impracti¬ 
cal. The chances for slant attack by interception ap¬ 
pear somewhat enhanced because, if the interceptor 
aims to intercept one airplane, he may miss it but 
catch another, but the interceptor would also risk 
drawing fire from several of the bombers. 

The vulnerability of a mass formation could un¬ 
doubtedly be reduced by providing special arma¬ 
ment and/or armor for the airplanes assigned the ex- 



MASS RAID TACTICS 


327 


treme positions in the formation at the expense of 
their bomb load. This has the design disadvantage of 
a duplicity of type and the tactical disadvantage of 
more replacement parts and limitations of the tacti¬ 
cal disposition of the units. However, if mass forma¬ 
tion raids are contemplated, it may be well worth 
while to create not only the full-load bomber but also 
a protective fighter having the same performance and 
range with reduced or even with no bomb loads. 

The fighter version would have some forward¬ 
firing guns, possibly capable of elevation, and pos¬ 
sibly armor for the occupants. The elevated guns 
need not be continuously movable in elevation. It 
may suffice to provide two elevation positions: for 
instance, (1) parallel to the high-speed flight path 
and (2) at 10 or 15 degrees elevation therefrom. 

In view of the probability of most interceptions’ 
winding up in a tail chase, the bombers—or at least 
those assigned the rear positions in the formations— 
might deserve rear armor. 

A 7 5 Formation Shape 

The shape of the mass formation has an influence 
upon the mutual assistance between units. At high 
speed, considerations of the prevalence of tail attack 
may call for different formation shapes than at 
speeds at which the interceptors can attack from all 
quarters. 

Lone stragglers behind would have to rely on their 
own fire power to ward off pursuit. 

A compact phalanx with many bombers abreast 
along the trailing edge of the formation has the ob¬ 
vious advantage of drawing any tail attacker into the 
defense fire of the adjacent bombers at good azi¬ 
muths and almost at the same range with almost no 
lead correction requirements. 

A dense packing of bombers in several tiers above 
each other (three-dimensional formation), all termi¬ 
nating in one huge vertical trailing plane, would fur¬ 
ther enhance mutual assistance against the tail chase 
into which any persistent interception must develop. 

Single leaders or navigators ahead of the main pha¬ 
lanx may not be particularly endangered, provided 
some of the bombers in the front line of the main 
formation are equipped with a forward-firing gun. 

If some such forward-firing guns are distributed 
throughout the formation, it might serve to discour¬ 
age any interceptors from trying to break into the 
formation from above or below in case any of them 
have enough extra ceiling or climb (or rocket boost). 

A 7 4 Formation Density 

The safest density of the formation will probably 
depend not only on the degree of mass formation 


training attainable under combat conditions but also 
very much upon the concentration of interceptors 
loosed against a mass raid. 

Obviously, any three-dimensional pattern with 
several bombers in tiers above each other would have 
to open up upon arrival at the objective in order to 
avoid hitting the lower ships with bombs from the 
upper ones. The technique of this maneuver would 
hardly differ much from that of similar situations of 
bombers at lesser speeds, so that no new problem 
seems to arise from a boost of airspeed or altitude. 

If the rear bombers are to derive tactical aid from 
the firing power of their formation neighbors, then 
their spacing must remain dense enough; it should be 
but a fraction of the dangerous firing range. For ex¬ 
ample, assuming the bomber’s tail armament covers 
a cone of 30 degrees semiapex angle (i.e., 30 degrees 
off the fuselage axis), a spacing of more than 200 yd 
would render the bomber powerless to assist his 
neighbor when the pursuer approaches within 350 yd. 

A very much more scattered formation, with sev¬ 
eral hundreds or thousands of yards spacing be¬ 
tween bombers, would cover a huge area. For in¬ 
stance, 400 bombers in a single-layer square lattice of 
2,000 yd spacing would cover 500 square miles of sur¬ 
face. This would be highly confusing to any ground 
organization trying to dispatch interceptors. How¬ 
ever, once the interceptor squadron encounters any 
part of the raiding force, a man-to-man duel is likely 
to result, with no further immediate advantage to 
the bomber; on the contrary, in order to prevent dis¬ 
organization of the raid schedule, the individual 
bomber will refrain from escape maneuvers and stay 
on his job. 

Mass raids in waves of bomber groups, in dense 
formation of each group, timed to arrive at the inter¬ 
ception gateways at intervals intended to exhaust the 
interceptor force will probably have even greater 
merits at ultra-high speeds than they have at current 
moderate speeds. If they are timed to arrive at inter¬ 
vals equivalent to the interceptors’ flight duration, 
then all the fighters sent up to intercept the first wave 
will be out of fuel and out of commission for the next 
alarm. Furthermore, while they land they are mak¬ 
ing so many airports less available for take-off of the 
next lot. It may even be a good trick to disperse the 
first wave into several spaced groups 50 or 100 miles 
apart in order to arouse interceptors from many 
places so as to disorganize many stations when the 
next, more concentrated wave comes through. Stra¬ 
tegic choice of targets for various waves may aid in 
creating further confusion of ground defense meas¬ 
ures. The opportunities for such long-range strategy 
increase rapidly with the speed and range of the craft 
as distances shrink and decoy and evasion detours be¬ 
come more feasible. They force the defender to spread 



328 


APPENDIX A 


his interception out thinner. Similarly, the long-range 
and high-speed bombers cooperating in a mass raid 
can be directed to disperse and feign repulsion—only 
to reunite, after the fuel supply of the interceptors 
has been exhausted, to complete the bombing raid 
and return to their bases. 

A8 PRELIMINARY CONSIDERATIONS 
OF EFFECTIVENESS OF FIRE 

The previous chapters of this investigation con¬ 
tain a great deal of evidence of the overwhelming pre¬ 
ponderance of tail combat phases to be expected 
when an ultra-high and fast bomber is to be inter¬ 
cepted by fighters having but a small speed advan¬ 
tage. In a qualitative way it appears obvious that 
any defensive devices that impair the bomber’s per¬ 
formance may work to its own disadvantage, rather 
than to its safety. The performance is impaired by 
the weight of armor and armament and, even more, 
by the drag of protruding parts, which increases rap¬ 
idly with the degree of angular gun coverage. A quan¬ 
titative treatment would require expression of the 
performance loss in terms of armament coverage. In 
view of the multiplicity of means (guns of various 
caliber, fixed mounts, eyeballs, and turrets), this is 
hardly feasible. However, it is obvious that arma¬ 
ment covering a limited tail cone and one or two 
fixed, essentially forward directions can be accom¬ 
modated at an almost negligible drag and an easily 
tolerable weight penalty, whereas full coverage of all 
angles would entail either many guns involving a 
very heavy weight penalty or turrets involving a pro¬ 
hibitive drag penalty. 

A 81 Limitations of Probability Calculation 

If it were possible to show that the chances of be¬ 
ing hit from certain angles are entirely negligible, then 
defense against attack from these angles would be 
wasteful on any count. If it were possible by rigorous 
methods to arrive at a quantitative probability dis¬ 
tribution of hits from various angles in space, then 
the decision of armament limitations of the bombers 
could be put on a scientific basis. The following study 
does not pretend to accomplish all this, but merely to 
analyze the factors governing the chances of hits in 
interception and to substantiate the inference that 
ultra-high performance is more valuable to the bom¬ 
ber than defense against ineffective combat phases. 

Relevant Influences 

The probability of a hit in interception from any 
particular angle off the tail of the bomber depends on 
many factors, such as: 


1. The speed of the bomber at the high altitude of 
the raid and his initial maneuvers. 

2. The speed of the interceptor, both in the climb 
and in high-altitude level flight, and the speed losses 
due to maneuvering. 

3. The probability of any one initial position at the 
instant when the interceptor detects, identifies, and 
picks his quarry. 

4. The mechanism of direction or the technique of 
the chase up to arrival at firing range. 

5. The load factors that the interceptor can or will 
tolerate. 

6. The maximum and minimum firing range. 

7. The technique of the chase while in firing range. 

8. The influence of the chase maneuver (range, 
banking, and acceleration) upon the interceptor’s 
aim. 

9. The perfection of the interceptor’s aiming de¬ 
vices, especially its correction for phoronomic in¬ 
fluences. 

10. The number and distribution of guns on the in¬ 
terceptor, their caliber, and firing speed. 

11. The interceptor’s firing tactics (burst density 
versus range, azimuth, bank, etc.) 

12. The influence of range on scatter and penetra¬ 
tion. 

13. The size and pattern of vital areas on the bomb¬ 
er and the influence of azimuth thereon. 

14. The chances of the interceptor to approach to 
any particular proximity and azimuth without being 
warded off by the bomber’s defense or by the fire 
from other bombers in the same formation. 

15. The influence of cooperation between several 
fighters attacking the same bomber upon items 4, 7, 
11, and 14. 

16. The effect of the bomber’s escape maneuvers 
upon items 4, 5, 7, 8, 11, 13, 14, and 15. 

A 8 3 . Assumptions 

It is at once apparent that the variety and multi¬ 
plicity of these influences alone constitute too com¬ 
plex an aggregate to cover completely. Only by the 
most radical process of elimination of variables and 
selecting “representative” examples shall we be able 
to reduce the problem to a useful digest. If this is at¬ 
tempted here at all, it is done with a keen realization 
of the fact that different assumptions will lead to dif¬ 
ferent results, but it is contended that the examples 
selected are representative in that they indicate a 
significant trend. 

1. Let us assume (a) the bomber’s speed v = 450 
mph at 40,000 ft and (b) that it is not diminished in 
combat and that he continues in straight level flight. 

2. Let us assume (a) the interceptor’s speed V = 
500 mph at 40,000 ft and (b) that it is not appreci- 



PRELIMINARY CONSIDERATIONS OF EFFECTIVENESS OF FIRE 


329 


ably diminished in combat. (Taking this speed loss 
into account will favor tail combat phases.) 

3. Let us arbitrarily assume that the range at 
which the interceptor detects and selects his victim 
is a definite value, say 5 miles or 8,800 yd. If the in¬ 
terceptor pilot’s knowledge of the whereabouts of the 
bomber before he detects and identifies it on his own 
detector is but approximate or inaccurate, he may 
patrol an assigned area at random course. However, 
this assumption might introduce a systematic error 
because the chances are that the pilot does have a 
fair idea of the bomber’s whereabouts and he already 
heads for him. The azimuth at which he is most likely 
to come upon his quarry will therefore depend upon 
the interception strategy. Not knowing whether this 
will be in the nature of a patrol barrier or of a long- 
range forward lunge or a mass melee after the in¬ 
vader has penetrated deeply into the territory, we 
may defend a random-course-at-detection assump¬ 
tion as representing a sort of grand average of various 
such strategies. 

With this assumption—namely, all initial course 
angles 0 O of the interceptor equally likely—the prob¬ 
ability for the bomber to be detected from any partic¬ 
ular azimuth a 0 is then 

<t>=irl 2 

cos <f> dd o 

-W 2 

with . 

, sin do 

<f>o = tan - ao (31) p 

-cos 6o 

e 

This integral has been computed for 1/e = 0.9 for 
azimuth steps of 10 degrees and the resulting proba¬ 
bility distribution plotted against azimuth a 0 in Fig¬ 
ure 31. 

4. It may be expected that, after detection and 
identification of the quarry, the pursuit takes a 
course somewhere between the scopodromic and bal- 
lodromic concept. (Let us disregard the alternative of 
brachydromic interception since it may be contended 
that its probability is very much smaller because of 
the greater skill required for its execution, in the 
very much shorter combat duration and its consider¬ 
ably lesser fire concentration.) In fact, the scopo¬ 
dromic technique is but a special case of ballodromic, 
namely, with \f/ = 0. 

In reality the interceptor will at first have to stick 
close to the scopodrome until he gets close enough to 
determine the relative motion of the quarry accur¬ 
ately. During the passage through the region of high 
load factor, even if he wanted to open fire already, he 
would have to allow for his high angle of attack, 
which puts him between the scopodrome and the 


p For derivation see Section A.8.6. 


theoretical (path-fixed gun) ballodrome. Later, he 
approaches the latter more closely, but then they 
differ but little, and the azimuth will have shrunk to 
very acute values. 

The probability of arriving at various firing ranges 
at various azimuths can be computed by transferring 
the initial probabilities along or between the scopo¬ 
dromic and ballodromic curves of Figures 2 and ll. q 
Figure 31 shows examples of the effect of scopodro¬ 
mic approaches from 5 miles detection range. 

5. When arriving at firing range, however, the ap¬ 
proach may be thwarted or hampered by high load 
factors. The tolerable load factor depends on the 
time of exposure. For simplicity’s sake, let us assume 
that a load factor of 5 g is the limit for the interceptor 
and that if his path would tend to lead to higher loads, 
he would ease up and drop into the sector between 
4 g and 5 g (more crowded toward 5</). The result of 
this restriction is also shown in Figure 31 in dotted 
lines. In reality, at least the later phases of the attack 
must be more nearly ballodromic in order to hit. The 
transition means a temporary increase of load factor 
followed by some relief. 

6. The probability of receiving effective fire from 
any particular azimuth depends very much on the 
firing range of the interceptor. The critical azimuth 
at which the maximum load factor is attained is of 
the order of 56 to 58 degrees and the critical range at 
this azimuth for 5 g is 550 yd (with u = 2,200 fps 
muzzle velocity) and 690 yd (with u = 2,875 fps). If 
the firing range begins at but little less than this, then 
the combat azimuth region is very narrow; as larger 
firing ranges than the critical are considered, the reg¬ 
ion of combat azimuths expands rapidly. 

As to the minimum range at which the engagement 
is considered decided or broken off, let us arbitrarily 
assume it to be 150 yd. Changing this ± 50 yd would 
not affect the result very much. 

7. Let us assume now that during combat the 
quasi-ballodromic chase continues, and the pursuer 
keeps creeping up without throttling until he arrives 
at a certain minimum range where he breaks off the 
engagement to avoid collision. (This assumption is 
rather arbitrary. It tends to enhance the probability 
of hits from close range and acute azimuth near the 
tail. Any departure of the interceptor from these tac¬ 
tics will affect the conclusions. If he throttles down to 
the same speed as his quarry, he can extend the dura¬ 
tion of any range convenient to him at will, but this 
will only crowd the probabilities of hits more closely 
around the tail a = 0.) 

8. The accuracy of aiming the airplane by the gun- 
sight at a distant target undoubtedly depends some- 


q The calculation is explained in Section A.8.6 for scopo¬ 
dromic approach. 








330 


APPENDIX A 



Figure 31. Azimuthal distribution of encounter probability at any range r\ in terms of the detection range r 0 . 


what on the size of the target as it stands out against 
its background, but sufficient information is not 
available to express this influence in mathematical 
terms. Therefore it will have to be left implied in a 
generous assumption of the influence of range upon 
scatter to be introduced separately. 

A transient surge of acceleration (load factor) has a 
very pronounced influence on aiming accuracy inso¬ 
far as it impairs the steadiness of the pilot-gunner’s 
head position and his precise coordination. It must 
broaden the scatter pattern in the direction of the re¬ 
sultant inertia force, perhaps somewhat in proportion 
to the excess of the load factor over unity. 

The banked attitude of the airplane causes the 
trajectory correction of a conventional fixed gunsight 
to be in the wrong plane. The error due to this in¬ 
fluence would tend to increase approximately in pro¬ 
portion to the square of the range and the sine of the 
banking angle. Not enough information is on hand to 
assign a definite value to the proportionality factor. 
As a compromise it may perhaps be arbitrarily as¬ 
sumed that the last two errors combined will tend to 
double the scatter pattern at 3 g and treble it at 5 g. 

9. If the gunsight of the interceptor is fixed so that 
he has to make allowance for leading his target by es¬ 
timate and keep his crosshair way ahead of the tar¬ 


get, large errors will be introduced whenever a large 
lead angle is required. However, even where the sight 
is equipped with a lead corrector to bring the cross¬ 
hair back on the target, the range and turning rate or 
the target speed and aspect have to be first deter¬ 
mined and then cammed into the sight. At best this 
introduces guessing errors which increase with range 
and with the sine of the azimuth. Again it is difficult 
to assign definite probability values to these errors 
beyond stating that they cause the average scatter 
area to increase with a. Practical experience may 
eventually be accumulated and applied to arrive at 
a more rational choice for this coefficient. Reference 
si made to item 13. 

10. The number and the firing speed of the guns de¬ 
termine the density of the fire but they will have no 
influence upon the relative probability of hits being 
attained from any one quarter, although they do have 
an influence upon the seriousness of transient com¬ 
bat phases, such as those discussed in the section on 
brachydromic passage. 

11. While the desire to conserve ammunition will 
prompt the interceptor to hold his fire as long as the 
defense of the bomber permits him, it may be as¬ 
sumed, (a) for simplicity’s sake and (b) in order to 
stay on the conservative side for the bomber, that 







PRELIMINARY CONSIDERATIONS OF EFFECTIVENESS OF FIRE 


331 


the firing is done continuously (or at regular uniform 
burst intervals) from a certain long firing range r e toa 
certain short firing range r $ . In reality, the tendency 
to fire more or longer bursts at shorter range would 
favor the acute tail azimuths. 

12. To arrive at any definite probability values, the 
primary scatter function which is characteristic for 
the gun and its installation would have to be known 
for various conditions. 

13. The size of the vital area is assumed to be ap¬ 
proximately 700 sq ft, which comprises either the cen¬ 
tral part of the fuselage in lateral aspect or the four 
engines and the fuel tanks in rear aspect. Neglecting 
the variation of the vital area with azimuth helps a 
great deal to simplify the problem. This simplifica¬ 
tion is, however, rather arbitrary, especially if it also 
neglects the difference in compactness of the vital 
area. In rear aspect the height of the vital area is 
smaller and its span larger than in lateral aspect. 
This works in a way to compensate for the variation 
of errors in trajectory drop—errors which are most 
serious at the longer ranges which are more likely to 
be associated with large azimuths. It is proposed, for 
simplicity’s sake until more accurate knowledge is 
gained, to offset this aspect influence against the 
influence of imperfect leading described in paragraph 9. 

14. For the present purpose, let us disregard the 
possible influence of defense fire from the bombers 
upon the attacker. One may later attempt to ap¬ 
praise the relative chances of the two craft. 

15. Let us merely assume that cooperation between 
interceptors concertedly attacking a single bomber is 
effective in diverting the latter’s fire so that as¬ 
sumption (14) is justified. 

16. Let us assume the bomber refrains from execut¬ 
ing any escape maneuvers and keeps on flying 
straight, relying solely on his own speed to let all at¬ 
tacks eventually develop into tail chases. 

A84 Examples 

Under these numerous and drastic assumptions, 
the relative probability of a hit in a vital area from 
any azimuth can be evaluated theoretically by com¬ 
putation and shown graphically. 

It is certain, however, that under any of the as¬ 
sumptions proposed, by far the most of the serious 
fire will come from acute azimuth angles near the tail. 

A85 Interpretation 

The likelihood of the interceptor’s approaching to 
a range of effective fire at acute azimuth in pursuit of 
a bomber flying straight will depend on the relative 
chances of the attacker and the attacked to put each 
other out of combat during the earlier phases of the 


approach. At any given azimuth, these chances are 
governed by the influences of range, target aspect, 
lead required, azimuth, bank and load factor, the 
ballistic air velocity on bullet impact, and armor. It 
may be justified to make the same assumptions as 
outlined in Section A.8.2 except that the interceptor 
is likely to be a smaller target, say 400 to 500 sq ft. 
But this may be approximately offset by the fact 
that a flexible gun of the bomber can fire at appreci¬ 
able azimuth, with negative lead correction fully cor¬ 
rected for by a simple mechanically compensated 
sight like the French Alkan sight; this correction is 
accurate enough for the assumed bomber’s high 
speed V, whereas the additional forward lead to be 
guessed to allow for the pursuer’s cross-field velocity 
component is only a negligible fraction of that re¬ 
quired by the interceptor’s gun. As to the ballistic 
advantage of the bomber, let us assume that it is a 
stand-off against the interceptor’s protected fuel 
tanks. 

The decisive difference between the two craft then 
remains the presence of the load factor on the pur¬ 
suer and its absence on the bomber. A quantitative 
appraisal of relative chances of the two combatants 
therefore depends largely on the assumption made in 
item 8 and on their relative fire power (number of 
guns in action). 

A 8 6 Probability of any Azimuthal Position 

The probability of arriving at a definite fixed de¬ 
tection range r 0 at any particular detection azimuth 
« 0 is assumed to be proportional to the cosine of the 
angle <t> at which the relative flight vector of the pur¬ 
suer intersects the range vector. If all course angles 
6 0 of the pursuer are equally likely, the probability 
per unit sector is p« 0 = 1/2;r J cos <f> dd 0 ; but of these 
only the positive values of cos <j> are to be taken as 
representing approaches. The negative values would 
represent recession from closer range, i.e., stepping 
out of the detection range circle rather than into it; 
they must be disregarded. The integral of all prob¬ 
abilities for all ao is therefore less than unity, mean¬ 
ing that many courses 0 O are misses. (See Figure 32.) 

To express 0 in terms of 6 0 , observe on the speed 
vector diagram that V - v = TF is the relative speed 
vector of the two craft making an angle co with pur- 
sued’s path. 

In the vector triangle the law of sines says 
e sin (co + 0 O ) = sin oj 

from which 



and 0 = 180 — u — otQ. 




332 


APPENDIX A 



In this way values of cos <f> were computed for 10- 
degree steps of course angle 0 O , plotted against 0 O for 
each a 0 ; their positive parts were then planimetered. 
The result of this procedure is the master graph in 
Figure 31 labeled “detection range.” 

A 9 MOCK INTERCEPTION TO SCALE 

Interception of aircraft at ultra-high speed and al¬ 
titude involves techniques which differ from those 
applicable to conventional interception practice in 
many respects. It is therefore appropriate to study 
the new techniques not only theoretically but also 
practically as far as possible in advance in order to be 
prepared to use the new craft to full advantage as 
soon as it becomes available. Any extrapolation from 
lower-speed to higher-speed operation must be gov¬ 
erned by certain scale rules lest it be misleading. By 
observation of proper scale rules, however, it should 
be possible to gain some insight into the practicality 
or difficulty of some of the maneuvers theoretically 
studied, into the practical validity of some of the as¬ 
sumptions made in the theoretical studies, and into 
the chances of interception combat in the various 
maneuvers investigated. 

A 91 Scale Rules 

As a general rule, in order to maintain geometrical 
similarity of the relative motion between the mock 
maneuver and its prototype, the ratio between the 
true airspeeds of the two craft must remain un¬ 
changed and the speed advantage of the pursuer 
must be the same percentage of his true speed in both 
cases. However, this will not always allow the load 
factor (acceleration) to remain the same in corre¬ 
sponding phases of those maneuvers which involve 
turning in any plane; only in straight-flight intercep¬ 
tion like the brachydromic passage do no load factor 
problems arise. 


A92 Instrumentation 

Aircraft to be used for mock-interception studies or 
practice should be equipped with accelerometers, gun 
cameras, gunsights, and airspeed meters on the dials 
of which the true airspeeds desired to be held are 
suitably marked with proper density correction for 
the altitude at which the maneuvers are to be ex¬ 
ecuted. On the bomber, if the defense is to be sim¬ 
ulated by a gun turret or flexible tail gun, some 
means should be provided to record the gun excur¬ 
sion, or at least its azimuth, on the gun camera film 
or on some separate recorder adapted to be syn¬ 
chronized or otherwise tied in with the gun camera. 

In planning the various maneuvers Figure 33 will 
be helpful. It shows the turning times for the comple¬ 
tion of a 180-degree nonskid horizontal turn at vari¬ 
ous constant speeds and properly banked so as to 
attain and hold certain g values on the accelerometer. 

It should be remembered that all course changes in 
mock maneuvers must be executed with reference to 
true compass points and not with reference to land¬ 
marks because otherwise wind might distort the 
maneuvers appreciably. 

Stretched Time Scale 

If, in simulating any curved approach at reduced 
speed scale, the distances between the two craft at 
various maneuver phases of the mock combat are 



Figure 33. Turning times for completion of horizontal 
turns. 






MOCK INTERCEPTION TO SCALE 


333 


correctly reproduced, the problems of target recog¬ 
nition, range determination, and aiming accuracy 
may appear in “full size” but the time available for 
the planning and execution of the maneuver is 
stretched in inverse proportion to the speed reduc¬ 
tion, and the severity and precision of the accom¬ 
panying banking operation are much reduced, ac¬ 
cording to Figure 33. 

For instance, in practicing scopodromic or ballo- 
dromic approaches of various parameters, i.e., from 
various abeam distances rr and also from various ini¬ 
tial azimuths a 0 while at a given initial range r 0 , say 
at speeds v = 333 mph pursuit vs 300 mph bomber, 
to simulate 500 and 450 mph respectively, at a speed 
scale of 2:3, each mock maneuver will take % as 
long as the real one, and a load factor corresponding 
to 3.42# will represent 5 g in the real case/ 

A 9 4 Reduced Range Scale 

It would also be interesting to study or interpret 
such maneuvers on the basis of time available rather 
than of correct distance. This simply amounts to in¬ 
terpreting a mock approach to represent a real one 
from % the distance at every azimuth. Then the 
target will appear 3 / 2 times too large at every phase 
but the load factors ( g ) and the banking are correctly 
represented, and a vivid picture of the difficulty of 
performing the maneuver and aiming during the 
banked phases is conveyed. 

A 9 5 Practice Maneuver Pattern 
for Curved Approach 

In executing specified mock approaches from defi¬ 
nite ranges, it may be difficult to determine these 
ranges in flight. Post mortem evaluation can prob¬ 
ably be reconstructed from the gun camera records. 
Besides, the following procedure may be helpful in 
planning the mock flights somewhat as depicted in 
Figure 34. 

First fly the two craft in close formation at a cer¬ 
tain (say 0 degrees) compass course to check their 
airspeed meters against each other. Then agree by 
radio on the speeds to be maintained during ma¬ 
neuvers. Upon a definite signal the pursuer turns 90 
degrees off the course (say to the right, course 90 
degrees). The bomber flies straight on for a prear¬ 
ranged number of seconds (say 30) and then com¬ 
pletes a 180 degree right turn carefully to maintain a 
definite average load factor (say 2.5) during the turn 

r Load factors in horizontal turns are not exactly propor¬ 
tional to the speed ratio v\ : Vi but 

rh/fh = VWHWH 5 ) 
where R is the turning radius. 


in order to complete it in a definite time (say 17 sec). 
He then proceeds straight and level at the same orig¬ 
inal altitude and speed in the opposite direction 
(course 180 degrees). The pursuer also turns at the 
same rate, either after the same lapse of time since 
the break-away or after a shorter or longer period ac¬ 
cording to a prearranged schedule; he turns in the 
opposite sense (say to the left) to maintain sym¬ 
metry. He may not have to complete a 180-degree 
turn but picks up his target, boosts his speed to the 
desired speed advantage, and heads for his quarry. If 



Figure 34. Practice maneuvers (S = scopodromic, 
Ba = ballodromic, Br = brachydromic). 


he turned simultaneously with the bomber, then he 
can sweep into a scopodrome after completing but 
135 degrees of a turn (to course 315 degrees), and he 
will begin the approach at 135 degrees bomber’s azi¬ 
muth. The later the pursuer turns, the greater is the 
parameter rr of the ensuing pursuit curve. If the pur¬ 
suer waits to begin the turn after twice the time the 
bomber flew away before he started his, then the dis¬ 
tance traveled straight by the pursuer is approxi¬ 
mately r T . The bomber will be found dead ahead on 
course 270 degrees, and the scopodrome will be large, 
winding up in a tail chase before entering firing range. 










334 


APPENDIX A 


The sooner the pursuer turns, the sharper and the 
more vicious the critical phase of the approach ma¬ 
neuvers becomes. 

It must be noted that the banking maneuver dif¬ 
fers in character for different speed ratios. If the pur¬ 
suer has an ample speed advantage, then the phase 
of steepest banking occurs very late, while already 
close to the quarry’s tail. If he flies twice as fast as 
the bomber, the pursuer has to bank more and more 
as he approaches collision. However, if the speed ad¬ 
vantage is but a small fraction of either craft’s speed, 
then the pursuer has to “unbank” again, viz., begin 
leveling off at a certain position in the chase. As the 
speed advantage diminishes, the critical phase of 
steepest banking approaches the position where the 
pursuer passes 60 degrees off the pursued’s tail. For a 
speed ratio 10:9 this critical azimuth is about 56 
degrees. As soon as the pursuer passes this azimuth, 
he has to give counter aileron and begin to unbank, 
even faster than at the rate he banked before. Other¬ 
wise he overshoots. 

Practice Maneuver Pattern 
for Straight Passage 

In simulating brachydromic passages, again either 
the ranges can be considered full scale and the time 
stretched as in a slow motion picture or else the time 
lapse can be reproduced correctly by reference to pro¬ 
portionally closer range phases. In practicing brachy¬ 
dromic passages, the trick for the pursuer to learn 
consists of the following maneuver. Aim ahead of the 


bomber so that the latter appears at an angle off your 
bow slightly less than that at which you would esti¬ 
mate your ship appears off his bow; if you find the 
course where this “lead” or relative crab angle does 
not change while you fly straight, it will lead to colli¬ 
sion (for safety’s sake the pursuer may fly at a 
slightly higher level than the bomber); during the ap¬ 
proach, correct your course slightly, so that the tar¬ 
get seems to creep up slowly to smaller and smaller 
lead angles and finally passes dead ahead at short 
range, all while both ships fly straight. 

As an aid in executing practice maneuvers of this 
sort, the two cooperating pilots may again start out 
flying in formation; then on a signal the pursuit 
breaks out at right angles and they both fly straight 
for a predetermined number of seconds (or until they 
signal each other by radio). Now both make a 180- 
degree turn in opposite sense and so that they head 
for interception near the original parting point, as 
depicted in Figure 34. Thus, only slight course cor¬ 
rection (toward the bomber) will be necessary for 
the interceptor to properly “miss” the bomber and to 
let him fly past ahead, while getting in a “burst” dur¬ 
ing the passage. 

A 9 7 ’ Escape Turns 

Maneuvers involving escape turns on the part of 
the bomber may be practiced either under exactly 
the same time stretch and acceleration-reduction 
consideration or else under the same range scale equals 
speed-scale condition as outlined in Section A.9.4. 



Appendix B 

LATERAL STABILITY OF HOMING GLIDE BOMBS 
WITH APPLICATION TO NAVY SWOD MARK 7 AND MARK 9 


B 1 LATERAL STABILITY EQUATIONS 

The motion of a glider in flight is determined by 
two factors: the aerodynamic forces and moments 
due to the reaction of the air on various parts of the 
glider, and the forces and moments due to gravity. 
The resultant of the gravitational forces may be rep¬ 
resented by a force equal to the weight of the glider 
acting vertically downward through the center of 
gravity. The resultant of the aerodynamic reaction 
is conveniently represented by three mutually per¬ 
pendicular forces acting through the center of gravity 
and three moments acting about these three mutu¬ 
ally perpendicular axes which meet at the center of 
gravity. In general, a glider has a plane of sym¬ 
metry which in normal steady flight includes the 
direction of motion. For convenience, we will choose 
axes as follows: the X-axis is taken in the plane of 
symmetry in the direction of the relative wind dur¬ 
ing the steady-flight condition, the Y-axis perpendic¬ 
ular to the plane of symmetry, and the Z-axis in the 
plane of symmetry and perpendicular to the X-axis. 
Rotation about the X-, Y-, and Z-axes are denoted 
by the angles </>, 0, and \p respectively, and angular 
rates of rotation about these axes by p, q, and r. 

We may ordinarily consider the motion of a glider 
as divided into two independent types of motion. One 
type, called longitudinal motion, includes motion 
that does not displace the plane of symmetry of the 
airplane, and the stability *of motion in this plane is 
termed longitudinal stability. 1 The other type of mo¬ 
tion, called lateral motion, includes all components 
that do displace the plane of symmetry, and the 
stability of this motion is termed lateral stability. 
The longitudinal stability of homing glide bombs is 
discussed in detail in reference (1), and will not be 
considered further here. The present discussion of 
lateral stability refers principally to gliders in which 
the aileron is the only lateral control surface, al¬ 
though the discussion of stability with fixed control 
surfaces applies equally well to gliders equipped with 
both rudder and ailerons. 

The stability characteristics of a glider are studied 
from the standpoint of the motion obtained from 
small displacements from a state of equilibrium. Un¬ 
der equilibrium conditions the glider flies in a straight 
line at constant speed with the plane of symmetry 
vertical—wings level. The resultant air reaction lies 
in the plane of symmetry and passes through the 


center of gravity. Let us define the following quan¬ 
tities : 

7 = angle between flight path and horizontal 
<j> = angle of roll 

\J/ = angle of yaw 

8 = angular displacement of ailerons 
W = weight of glider 

m = W/g, mass of glider 

V = velocity along X-axis 

v = velocity in Y-direction (sideslip velocity) 
p = air density 
S = wing area 
b = wing span 

Y = force along Y-axis (lateral force) 

L = moment about X-axis (rolling moment) 

N = moment about Z-axis (yawing moment) 

p = ^ = rate of roll 

d\p e 

r = = ra ^ e °* ^ aW 

A = moment of inertia of glider about X-axis 
C = moment of inertia of glider about Z-axis. 

Let us also define the following coefficients: 


Lift 

Drag 

i pSV* 

L 

c - N 

h P SV 2 b 

w ipSV 2 b 

Y 

o V 

i pSV* 

V 


At equilibrium, the quantities L, N, Y, p, r, <t >, \p, 
and v are all equal to zero. 

Let us now assume a small displacement from 
equilibrium and determine the equations which gov¬ 
ern the motion. In order to reduce the complexity of 
the problem to enable a solution to be obtained with¬ 
out a prohibitive amount of calculation, the following- 
assumptions are made: 

1. Forces and moments on lifting surfaces are 
assumed proportional to the square of the airspeed. 

2. Forces are unaffected by angular velocities and 
angular accelerations, and moments by angular ac¬ 
celerations. 

3. The glider is assumed sjunmetrical, and thus 
lateral motions and longitudinal motions are as¬ 
sumed to be independent. 


335 





336 


APPENDIX B 


4 . The combined effect of two or more forces or 
moments is assumed proportional to the algebraic 
sum of the separate components. 

5 . The changes in aerodynamic forces and mo¬ 
ments due to a deviation are assumed proportional to 
the deviation. 

6 . Secondary effects involving the product of two 
small quantities are neglected. 

7. The principal axes of inertia of the glider are 
assumed coincident with the reference axes. 

The lateral equations of motion are given by: 


dv 17 .d\l/ 

m di + mV Tt 
dY 

= —j^v + mg sin (f> cos y + mg sin i p sin y ( 2 ) 


. d 2 <t> dL . dL . dL . dL . 

A-,— = -—v + p + -r—r -f —~8 
dt 2 dv dp dr dd 

f,d 2 \l/ dN . dN dN . dN 

C—L. = —-v + -—p + —r + —5 

dt 2 dv dp dr db 


(3) 

(4) 


Let us define the following quantities: 


1 

dL 

N p 

1 

_dN 

A 

dp 

c 

dp 

1 

dL 

N r 

1 

dN 

A 

dr 

c 

dr 

1 

dL 

N v 

1 

dN 

A 

dv 

c 

dv 

1 

dL 

N s 

1 

dN 

A 

d8 

c 

d8 


1 

dY 




<8 

II 

31 

dv 




(5) 


Remembering that <j> and \p are assumed to be 
small quantities, equations (2 , (3), and (4) become: 

^ ~ V-Jj + Y v v + (g cos y)<l> + ( g sin y)\p ( 6 ) 

W = L* + Lp ft + Lr ft + LsS (7) 

*±- N * + N » + N <!t + N * (8) 

Let us assume that v, <f>, and 8 vary with time 
according to laws of the form: 


V = V 0 6 Xt </> = <f) 0 € Xt 

t = \poe xt 8 = 8 0 e xt ( 9 ) 


and determine what values of X will satisfy the above 
equations. The stability of the various motions will 
be determined by the nature of the values of X that 
satisfy these equations. If X is positive, small dis¬ 
placements in these quantities will continuously in¬ 
crease with time, and the motion will be unstable. If 
X is negative, small displacements will decrease with 


the time, and the motion will be stable. If X is com¬ 
plex, the motions will be oscillatory, with increasing 
or decreasing amplitude depending upon whether the 
real part of X is positive or negative. 

If we substitute the values of v, <f>, and 8 given 
by equation (9) into equations ( 6 ), (7), and ( 8 ), we 
obtain: 

(X - Y v )v - (g cos y)<f> + (\V - g sin y)\J/ = 0 (10) 
-L v v + (X 2 - \L p )<f> - \L r t - U8 = 0 (11) 
-N v v - \N p <t> + (X 2 - XAW - N s 8 = 0 (12) 

Here we have three equations involving four vari¬ 
ables, since so far we have made no mention of the 
variation of 8 with the time. 


B2 STABILITY 

WITH FIXED CONTROL SURFACES 


If w T e assume fixed control surfaces, we set 8 equal 
to zero, and we have three equations in three vari¬ 
ables. To determine what values of X satisfy the 
above equations, it is necessary to determine only 
what values of X make the following determinant 


vanish. 

X - Y v 

— g cos 7 

XF — g sin y 

Li) 

X 2 - \L P 

— \L r 

—N v 

—\N P 

X 2 - \N r 


Solving the above determinant, we obtain the fol¬ 
lowing equation for X: 

X 5 + (-Nr - L p - FJX 4 

+ (LpNr - L r N p + L P Y V + NrY v + VN v )\ 3 

+ ( — L p N r Y v +L r N p Y v — L v g cos y — N v g sin y 

+ VL v N p — VLpN v )\ 2 +• ( L v N r g cos y 

— L r N v g cos 7 — L v N p g sin y 

-b LpN v g sin y)\ = 0 (14) 


This equation may be written in the form: 

AX 5 + B\ 4 + CX 3 + DX 2 + EX + F = 0 (15) 


where 


A = 1 


B = -L p - N r - Y v 
C = L p N r - L r N p + Y V (L P + N r ) + VN v 
D — ( L r Np — LpNr) Y v — L v g cos y 

— N v g sin y + V(L v N p — LpN v ) 
E = ( L v N r — L r N v )g cos y + ( L P N V 

— L v N p )g sin y 


F = 0 


(16) 


The determination of the values of X which satisfy 
equation (15) depends in general upon the solution of 
a fifth-degree equation. Since in this case, one root is 







STABILITY WITH FIXED CONTROL SURFACES 


337 


zero, it reduces to a fourth-degree equation. The 
values of the above quantities for a conventional type 
glider are such that the roots of equation (15) are one 
pair of conjugate complex roots, two negative roots, 
and, of course, one zero root. In general, the real 
parts of all the roots will be negative, corresponding 
to a stable condition, if all coefficients are positive 
and Rouths’ discriminant {BCD — D 2 — B 2 E) is 
positive. 

Let us now investigate which are the important 
terms in the equation in X and thus show how the 
nature of the solutions depends upon the values of 
the aerodynamic coefficients involved. In C , the 
quantity VN v is large compared with the other terms 
present, and in D, the quantity V{L v N p — L P N V ) is 
large compared with other terms. Let us, then, for 
the present, neglect the other quantities involved in 
these terms, and let C and D be given as follows: 

C = VN v D = V(L V N P - L P N V ) (17) 

In general, C and D will be large compared with B 
and E. Let us assume then that we may approxi¬ 
mately reduce equation (15) to the following form: 

[ X2 + ( B ~c) X + C ] [ X + |] [ x + £] x = ° 

(to; 

Multiplying, we obtain: 

x5+ ( B + |)x 4 + [c + ^(s-^) + f]x3 
+ {d + |[c + |(b-|)]}x 2 + £X = 0 (19) 


Thus, in order that equation (15) may be approxi¬ 
mately represented by an equation of form (18), the 
following relations must hold: 



In general, the values of the aerodynamic coeffi¬ 
cients for normal flight conditions of aircraft type 
missiles are such that these conditions are satisfied. 
Inserting the values of B, C, D , and E in equation 
(18) and assuming the approximate values of C and 
D given by equation (17), we obtain: 



O -'-' + tH '- 0 


( 21 ) 


With these approximations, X is given by: 
L V N P 


Xi — L P 
X 2 = - 


N v 
g cos t 


X3,4 


N r + Y v + 


± i 

X 5 = 0 


i _ 

i]/'vN v - ^ 


/ L v Nr - L r N v \ g sin 
\L V N P - L p Nj V 

• ) 


LvNp 

N v 


i Nr +Y V + 


L V N P \ 2 
N 


u ) 


( 22 ) 


Thus, v, <f>, and ^ can be represented by equations 
of the form 

v,<t>,t = Cl exp [” L p - 

n r 9 cos y(L,N, — L r N„\ , g sin 7 "] 

C 2 exp L y Vl„A T p -L p nJ + V J 

N7 )] 


+ 


+ C 3 exp [i(lM r + Y v + ^f 


(23) 


'Nr + Y v + 


L V N. 

N 


+ C, 


S=)- + e.] 


where C 1 , C 2 , C 3 , C 4 , and C 5 are arbitrary constants 
which have values depending upon which of the 
quantities v, <f>, and \J/ is being represented and the 
particular boundary conditions for that quantity and 
where exp R = e R . 

In general, the first term in equation (23) repre¬ 
sents a rapid subsidence, whose rate is a function 
mainly of L p , the roll damping term. This is the most 
important term in roll motion. 

The second term is a slow subsidence or diver¬ 
gence, whose rate depends primarily upon the magni¬ 
tude of L v N r compared with L r N v , and upon the glide 
angle 7 . This term determines the spiral stability 
characteristics of the glider; if the exponent is neg¬ 
ative, the glider is spirally stable; if positive, spirally 
unstable. In general, L v N r is larger than L r N v , and 
spiral stability is obtained, although the stability is 
always small. As seen from the second term of equa¬ 
tion (23), spiral stability is increased greatly at steep 
angles of descent (sin 7 — 1 ). 

The third term represents an oscillation whose 
period is determined largely by the term VN v , and is 
given approximately by T = 2tt/\/VN v . As is to be 
expected, this period depends mainly on N v , the yaw¬ 
ing moment due to sideslip, and the damping depends 
mainly on the term N r , the yawing moment due to 
rate of yaw. These are the main terms in the yaw 
motion. The constant C 5 occurs because of the fact 
that in a nonhoming missile, the zero point for meas- 



















338 


APPENDIX B 


uring angle of yaw is arbitrary; there is no preferred 
direction in space. 


B3 STABILITY OF SWOD MARK 12 
AND MARK 13 AIR STABILIZERS 


The following is a table of values of the lateral co¬ 
efficients applicable to the SWOD gliders: 


IF (lbs) 

S (sq ft) 
b (ft) 

A (lb sec 2 ft) 
C (lb sec 2 ft) 


dCi 



dCi 


dP 

dC n 



dC n 


dp 

dCy 

dp 


Mark 12 

Mark 13 

900 

1500 

18.3 

24.4 

8.4 

10 

23 

45 

75 

200 

-0.3 



0.2 C L 

-0.17 

-0.03C l 

- 0.2 

0.17 

- 0.6 


The values of IF, S, b, A , and C are measured 
values determined at the National Bureau of Stand¬ 
ards. The values of dCJdfi, dC n /dp, dC v /dp given in 
the table were obtained from a wind-tunnel test at 
the California Institute of Technology of an early 
model glider somewhat similar to the Mark 12 
glider 2 . However, since there are some marked differ¬ 
ences between this model and the present Mark 12 
and Mark 13 gliders, these values should be consid¬ 
ered as only approximate. Page 20 , Table 3, of refer¬ 
ence ( 2 ), gives the following values: 

Jf = 0 003 % = - 0 003 (24) 

and from Figure 12 of reference ( 2 ), 


dCc 

d+ 


= 0.010 


where C c is the crosswind force coefficient, which for 
small angles of yaw may be considered equal to C v , 


the lateral force coefficient. The values of the damp¬ 
ing terms 

dC t dCi dC n dC n 

•(£)’-(£)’*(£)’“ iw) 

are estimated by the methods described in 
references (3) and (4) and should be considered as 
only very approximate. 

Let us substitute the above values into the stabil¬ 
ity equations for a typical condition of flight of Mark 
13, with fixed controls in pitch. Assume Cl = 0.3, 
which corresponds approximately to an average posi¬ 
tion of the elevons of 10 degrees up from neutral. At 
equilibrium this gives a velocity of flight at sea level 
of V = 415 fps. 

Under this condition of flight, the following quan- 


tities 

have the values listed below: 

V 

= 

415 fps 



7 

= 


-12.5 degrees 


T 



dCi 

IpSVb 2 

= —4.0 sec -1 

Lj p 



/ x 

A 



°\ 

K2V ) 






dCi 

ipSVb 2 

= +0.8 sec -1 

lj r 


a 1 

( rb 

A 




l2F ) 



T. 



dCi 

ipSVb 

= —0.5 ft -1 sec -1 

lJv 



80 

A 

N p 



8C n 

IpSVb 2 

= —0.03 sec -1 



f P b \ 

c 




<2V ) 



N r 



8C„ 

ipSVb 2 

= —0.6 sec -1 



(J*_\ 

c 



d \ 

<2Vj 



N v 



8C n 

hpSVb 

= +0.1 ft -1 sec -1 



dp 

c 

y v 



6Cy 

ipSF 

= -0.15 sec" 1 



dp 

m 


Let us now substitute these values into equations 
(16), and thus determine the coefficients in equation 
(15) for X. We thus obtain the following equations 
for X: 

X 5 + 4.75X 4 + 44.6X 3 + 189.0X 2 + 9.81X = 0 (26) 

Xi = -4.35 X 2 = -0.053 

^ 3,4 = —0.172 + 6.55i X 5 = 0 (27) 

Xi represents a rapid subsidence which corresponds 
to the high damping in roll. X 2 is a slow subsidence 
determining the combined roll and yaw spiral mo¬ 
tion. X 3 and X 4 are a pair of conjugate complex roots 
which represent a damped oscillation in yaw. 






























LATERAL CONTROL SYSTEM OF SWOD MARK 7 AND MARK 9 


339 


If we put the values of the constants in equation 
(25) in the simplified expressions given by equation 
(21), we obtain: 

(X + 4.15)(\ + 0.057) (X 2 + 0.6X + 41.5) = 0 (28) 

Xi = -4.15, X 2 = -0.057 

X 8t4 = -0.30 ± 6.44* (29) 

These values are a fairly good approximation to 
those given by equation (27), except for the damping 
of the yaw oscillation. 


B 4 LATERAL CONTROL SYSTEM 34 
OF SWOD MARK 7 AND MARK 9 

In the case of homing gliders, that is, the case 
where the control surfaces are moved in such a way 
as to direct the glider toward a target, the manner in 
which the control surfaces are caused to move in 
response to the homing signals depends upon the 
characteristics of the particular servomechanism 
used. It is not possible to compute the effect of a 
general functional relation and to consider all possible 
specific relations which have been used as a basis for 
servomechanisms. We will consider here only the case 
of the lateral control system used in SWOD Mark 7 
and Mark 9. 5,6 The complete theory, taking into 
account the off-on link between the gyro and the 
servo, time lags in the servo, time lags in the homing 
control, and the complete set of lateral stability 
equations, would be a very unwieldy calculation, and 
it is thought that by treating these various factors 
separately as to their effects, the discussion may be 
made clearer. 

Lateral stabilization is obtained through the turn 
gyro, which is essentially a rate gyro equipped with 
electromagnet coils and electrical contacts. The elec¬ 
tromagnets are connected to apply torques to the 
gimbal frame which are proportional to the error 
angle in yaw as obtained by the homing device. The 
electrical contacts are arranged on opposite sides of 
the gimbal frame so that one contact or the other is 
closed, depending upon the sign of the sum of the 
torque applied by the electromagnets and the torque 
due to precession of the gyro wheel. Closing of a con¬ 
tact causes the servo to move the ailerons. The gyro 
is mounted in the gli ier at an angle so as to be sensi¬ 
tive to both roll and yaw. 

It was assumed in the derivation of the equations 
of motion that the principal axes of inertia of the 
glider were coincident with the reference axes. In gen¬ 
eral, this is not true, for in the case of SWOD Mark 7 
and Mark 9, the roll axis under normal equilibrium 
flight conditions is inclined to the direction of the 
relative wind by about 3 degrees. This difference has 


only a very small effect on the stability calculations 
for free flight, but must be taken into account in the 
case of homing flight. 

Since the flight path is on the average toward the 
target on a true homing course, the error angle as 
measured by the homing device will differ from the 
error angle referred to coordinates where the X-axis 
is in the direction of the axis of roll by an amount 
given for small angles by the following equation: 

r = t - P<t> (30) 

where y is the error determined by the homing de¬ 
vice, is the error angle, referred to coordinates with 
X-axis along the roll axis, /3 is the angle between the 
roll axis and a line from the glider to the target, and (f> 
is the angle of bank. 

Let a represent the angle the axis of the gyro 
makes with the axis of roll. Let c represent the rate of 
yaw in degrees per second that produces the same 
torque on the gimbal frame as an angular error of one 
degree. The particular contact on the gyro which is 
closed depends on the sign of the quantity co defined 
by 

a? = ~ + tan a ^ + c(^ - P<f>) (31) 

dt dt 


When co is positive, that contact on the gyro will be 
closed that causes the servo-control unit to move the 
ailerons differentially at a constant speed to produce 
a rolling moment to the left; when co is negative, the 
other contact will be closed, causing the ailerons to 
move at constant speed to give a rolling moment to 
the right. Thus the movement of the ailerons is al¬ 
ways in such a direction as to reduce the value of co to 
zero. A hunting motion is set up, the gyro contacts 
alternately closing and the ailerons moving alternate¬ 
ly for right and left differential. 

If we neglect this hunting motion, and assume that 
co is, on the average, zero, we have 

tan a + cxf/ — j 6c<f> = 0 (32) 

dt dt 

If we neglect the sideslip motion of the glider and 
assume, for the moment, that equilibrium of forces 
always exists along the lateral axis, equation (6) 
reduces to: 

0 = - V ^ + (g cos y )<t> + (g sin y)\p (33) 
dt 


In general, y is sufficiently small so that the term 
involving sin y may be neglected in equation (33) 
and cos y set equal to unity. Equation (33) thus 
becomes: 


d\p 

dt 



( 34 ) 



340 


APPENDIX B 


Substituting equation (33) and its time derivative 
in equation (32), we obtain: 


dV , 1 ( g cos y 

dt 2 dt tan a V V 


g sin y tan a 
V 



+ (l_+fftan7)c g _c^7 (35) 

V tan a 


If we use the simplified equation (34), which neg¬ 
lects terms in sin y and tan y, and assume cos y 
equal to unity, we obtain: 


dV , d±( _ g _ 

dt 2 dt\V tan a 


(3c \ 
tan a) 


+ 


cgt 

V tan a 


= 0 (36) 


This equation has the following approximate 
solution: 

* - *• j«a> [-?3(v - '*')]( • 

“(I4t5i;+') < 37 > 

This equation represents a damped oscillation in yaw. 

It is seen that the damping is influenced consider¬ 
ably by the value of /3c. The effect of this quantity, in 
general, is to reduce the damping. If the line from the 
glider to the target is below the roll axis (/3 positive), 
the damping of the oscillation is decreased; if it is 
above the roll axis, the damping is increased. As has 
already been noted, the average value of /3 for SWOD 
Mark 9 is about 3 degrees. /3 occurs in the damping 
term multiplied by c, the ratio of the rate of turn of 
the gyro to the error angle. The effect of /3 on the 
damping is thus emphasized by a large value of c. In 
the case of high sensitivity homing information— 
that is, a large signal for a small error angle—the 
damping may become negative, the oscillations be¬ 
come undamped and increase in amplitude to a limit 
where the homing information saturates on each 
oscillation, and these assumed equations no longer 
hold. 

Let us assume typical values of the constants in 
this equation. Let a = 20 degrees, /3 = 0.05, 
c = 0.5, and V = 415 fps. We obtain: 


t = to exp ( — 0.1030 cos (0.326£ + a) (38) 

This shows that an oscillation in yaw should occur 
with a period of about 19 se'conds and damping to 1/e 
of its initial value in about 10 seconds. 

Although this simplified theory predicts quite 
closely the period of the oscillation in yaw actually 
obtained in flight, the predicted damping is not ob¬ 
tained, and sustained hunting of about this period 
and of an amplitude of about 4 degrees is obtained. 
In order to increase this damping in yaw sufficiently, 
a bias gyro has been added to the control systems of 


SWOD Mark 7 and Mark 9. This gyro 5 is similar to 
the turn gyro except that it does not contain the elec¬ 
tromagnets, but only contacts, one or the other of 
which closes, depending upon the sense of the rate of 
turn. Its design and operation are described in detail 
in reference (5). This gyro is mounted in the glider 
along the average roll axis so that it will be sensitive 
chiefly to the yawing motion of the glider. Rolling 
motion will affect it slightly, because the axis of roll 
does not remain fixed for all conditions of flight. 

The bias gyro is connected in the circuit of the turn 
gyro so that when the rate of yaw of the glider is to 
the right, the right coil in the turn gyro is shunted by 
a resistor, and when the rate of yaw is to the left, the 
left coil in the turn gyro is shunted. The effect of the 
shunting current in the coils is shown in Figure 11 
of reference (5). It is seen that a bias is given to the 
signals put into the electromagnets in such a direc¬ 
tion as to oppose the yawing motion of the glider and 
thus increases the damping of the yaw oscillations. 
Experimentally, it has been found necessary to use 
this bias gyro to obtain sufficient damping in yaw. 

Figure 1 shows a typical flight of SWOD Mark 9. 
The curve labeled “Apparent Angular Horizontal 
Motion of Reflector” shows that an oscillation in 
yaw of period of about 20 seconds is evident, but it is 
of very small amplitude. 

Obviously the reason the damping predicted by 
the simplified theory is not obtained is due to the 
approximations made, which, in effect, neglect the 
effects of sideslip velocity, the time lags present in 
the homing signals, and the roll hunting motion in¬ 
volving time lags in the gyro and servo system. 

Let us now consider the effect of time lag in the 
homing intelligence. The homing intelligence has a 
time lag which is equivalent to that produced in an 
RC circuit of time constant of approximately 0.3 sec¬ 
ond. If we assume that the homing information has a 
lag corresponding to an RC circuit, equation (32) 
becomes 


dt , dj>, 

-37 + 37 tan a + c exp 
dt dt 


t f it ~ /3c) 

RC J RC 

ex P ^3 dt = 0 (39) 


Combining this equation with equation (34), we 
obtain: 

dV , ( g , 1 Vv 

dt 3 \V tan RC Jdt 2 

. ( 9 _ Pc \ dt 

\VRC tan a RC tan a) dt 


,_ cgt _ = 

VRC tan a 


(40) 


If RC is zero, this reduces, of course, to equation (36). 

















LATERAL CONTROL SYSTEM OF SWOP MARK 7 AND MARK 9 


341 


10 

5 

0 

-5 


o -10 


ELEVON DIFFERENTIAL 




co 

UJ 

cc 

UJ 

CL 


10 

8 

6 

4 

2 

0 


BIAS GYRO R 
5 


10 15 20 25 30 35 40 45 50 55 60 65 68 

RIGHT METER 




^ 


^.g. 20 25_ 30 35 40 45 50 55 60 65 68 


co 

UJ 

UJ 

cr 

o 

UJ 

o 


0 

-5 


1 

Iright 

-K- A A 1 

vw\ 

AuAaA 


1/WV 

i/\ 0 . 



/iAA/ 



v^VV-. 

V 

ii!CTvV\ 

/v wV 


V vv' 

AA/v 


^vy 

A7V\a 

VWV 

^/v V 





















CO 


o 

> 












-( z 

ERO POSITION 

ARBITA 

RY) — 



10 15 20 25 30 35 40 45 50 55 60 65 68 

APPROXIMATE DISTANCE FROM PLANE OF REFLECTOR (10^ FEET) 


- 1-1 -h 


H-1-1-h 


H- 1 - 1 -h 


H-1-1—I—I—h 


H-1-1—1 


24 23 22 21 20 19 18 17 16 15 14 13 12 11 10 9 8 7 6 5 4 3 2 1 0 


co 

UJ 


50 

40 

30 

20 

10 

0 


APPROXIMATE DISTANCE FROM REFLECTOR 


10 15 20 25 30 35 40 45 50 55 60 65 68 

A.G.C. METER 




Figure 1 Typical flight of SWOD Mark 9 








































































































































































































































































































342 


APPENDIX B 


If we substitute the values of /3, c, a , and V used in 
equation (38), and in addition let RC = 0.3, equa¬ 
tion (40) becomes: 

^ + 3.55^+ 0 .687f+°.355* = 0 (41) 

Solving, we obtain: 

^ exp - 3.38* 

4 - ^2 exp — 0.086£ cos (0.324£ + <r) (42) 

It is seen that the period of the oscillations is virtu¬ 
ally unchanged, but that the damping is decreased 
about 20 per cent. Sideslip effects and the hunting in 
roll with its various time lags involved are still 
neglected. 

To investigate the effect of sideslip velocity, let us 
consider the general stability equations, with an ideal 
servomechanism, where the differential of the ailer¬ 
ons 8, instead of increasing or decreasing at a con¬ 
stant rate depending upon the sign of co, will actually 
be proportional to co. Thus let us write 

8 = — Kco = + ^tan a + c\p — ^c<^j (43) 

If we assume 8 is given by an expression of the type 
8 = 8i exp (\t), we may write equation (43) as 
follows: 

8 + K(\ tan a - /3c) + K(\ + c)* = 0 (44) 

If this equation is combined with equations (10), 
(11), and (12), the values of X which satisfy the four 
equations are those that make the following deter¬ 
minant vanish. 


X 

- Y v 

— g cos 7 

XF — g sin 7 

0 



Lv 

X 2 - XL P 

— XL r 

-L8 

=0 


■N v 

-\N P 

X 2 - \N r 

-N8 


0 


K (X tan a — fic) 

K (X 4" c) 

1 

(45) 


When the above determinant is solved, an equa¬ 
tion of the fifth degree in X with 55 terms is obtained. 
This may be written in the form given by equation 
(15) with coefficients given in the following table: 

A = 1 

B = —N r — Lp — Y v -f- K[Ns 4~ Ls tan a] 

C = LpN r - L r Np + Y V (L P + Nr) + VN v 

+ KL S [—N r tan a — /3c + N p — Y v tan a] 

4- KNs [L r tan a — L p + c — Y v ] 

L> = Y v ( — LpN r + L r N p ) — L v g cos y 

- N v g sin y + V(L v N p — LpN v ) 

4- KLi[PcN r + cN p + N r Y v tan a + 0cY v 

- N P Y V + VN v tan a] + KN s [-/3cL r 

- cL p — L r Y v tan ot + L P Y V — cY v 
— VL v tan a] 


E = ( L v N r — L r N v )g cos y + { — L V N P 

+ L p N v )g sin y + KL d [—ficN r Y v — cN p Y v 
-b N v g cos y — ficVN v — ( N v tan a)g sin 7 ] 

+ KN s [PcL r Y v + cLpY v — L v g cos 7 
+ / 3cVL v 4- ( L v tan a)g sin 7 ] 

F = KL s [cN v g cos 7 4- (3cN v g sin 7 ] 

4- KN s [—cL v g cos 7 — (5cL v g sin 7 ] (46) 

Let us now substitute values for the coefficients in 
this equation for the typical flight conditions given 
by equation (25). We need values for the additional 
quantities La, Ns, K , a, and /3. From wind-tunnel 
tests 2 we may take dC e /d8 = 0.072, which gives 
L 5 = 80. We will assume for this flight condition 
that no yaw moments are produced by the ailerons; 
that is, Ns = 0 . Let K = 5, a = 20 degrees, and 
= 0.05, which are typical values. The equation for 
X becomes: 

X 5 4- 150.4X 4 4- 138. IX 3 4- 6,229X 2 + 953X 

+ 622 = 0 (47) 


This equation has the following approximate roots: 


Xi = —149.8, 

X 2i3 = -0.225 ± 6.43;, 

X 4 ,5 = -0.0765 ± 0.316t (48) 


Comparing this result with the free-flight results, 
equation (27), we see that instead of two subsidences, 
a damped oscillation, and a zero root, we now have a 
very rapid subsidence and two damped oscillations. 
The roots X 2 and X 3 represent a natural yaw oscilla¬ 
tion which is similar to that for the free-flight condi¬ 
tion and whose period and damping are mainly a 
function of the aerodynamic constants. The roots X 4 
and X 5 represent a damped oscillation whose period 
and damping depend primarily upon the constants a 
and c assumed for the ideal servo system. The period 
of this oscillation is about the same and the damping 
somewhat less than that obtained from equation (38), 
in which sideslip velocity was neglected. While the 
effect of sideslip is sometimes taken into account by 
introducing an “aerodynamic lag” in the simplified 
equation, it is important to note that there is no evi¬ 
dence of a true aerodynamic lag when the complete 
equations of motion are used. 

By examination of the solutions of equations (36), 
(40), or (45), it is seen that the damping in yaw 
should increase as a decreases. If we let a = 0 in 
equation (36), we obtain: 


d\f/ cxf/ 

dt + ' &cV 
9 


= 0 


Solving for \f/, we obtain: 

t = ^0 exp 


1 - 



f49) 


( 50 ) 








ROLL STABILIZATION SYSTEM 


343 


This equation shows that the yaw oscillation 
should reduce to a subsidence when a = 0. However, 
if we let a = 0 in equation (46), which takes into 
account sideslip velocity effects, we obtain the fol¬ 
lowing equation for X: 

X 5 + 4.75X 4 + 22.6X 3 + 173.7X 2 + 851X 

+ 622 = 0 (51) 

which has the following roots: 

Xi = -0.871 
X 2 ,3 = -4.08 ± 2.49f 
X 4> 5 = +2.14 ± 5.16i (52) 

We obtain the rapid subsidence and the damped 
natural yaw oscillation as in the case a = 20 degrees, 
but now X 4 and X 5 represent a rapidly divergent oscil¬ 
lation, and thus an unstable condition results. At 
some small value of a, the real parts of the roots X 4 
and X 5 change from negative to positive, changing the 
oscillation from a damped one to a divergent one. 


zero, and at t = 44, it is equal to its value at t = 0 . 
Thus a periodic hunting in the aileron motion is ob¬ 
tained of saw-toothed waveform, and with a period 
44. The aileron differential 8 as a function of the time 
is shown in Figure 2. 



Figure 2. Aileron differential 8 as function of time. 


B 5 ROLL STABILIZATION SYSTEM 

In order to study the effects of the roll hunting 
motion on the lateral stability, a more detailed dis¬ 
cussion of the roll stabilization system will be given, 
taking into account the off-on character of the link 
between the gyros and the servo clutches, and the 
time lag in the response of the servo. Let us assume 
that the glider is flying straight and level, that the 
hunting motion in roll has reached a steady-state 
condition, and that the motion is periodic. If the 
effects of rolling moment due to rate of yaw and 
rolling moments due to sideslip velocity are regu¬ 
lated, equation ( 7 ) for the motion in roll becomes: 

d S = L ”ft + L,s (53) 

At time t equals zero, let the rate of roll p equal p h 
the angle of bank <f> equal <f> i, and the rolling accelera¬ 
tion produced by the ailerons equal L&(d8 / dt)t\. Let 
us define K as equal to L 5 (d8/dt, and thus K becomes 
the acceleration in roil produced by the amount of 
differential on the ailerons developed by the servo in 
moving the ailerons at constant speed db/dt for one 
second. 

At the instant t = 0, the ailerons start to move 
differentially with a constant speed and in such a 
direction as to reduce the rolling acceleration. At 
time t = 4 , it is easily seen that the rolling accelera¬ 
tion will become zero, and at t — 24, it will become 
equal in magnitude and opposite in sign to its value 
at t = 0. At t = 24, it is assumed that the direction 
of motion of the ailerons reverses, so that at time 
t = 34 , the rolling acceleration is again reduced to 


We may express the motion in roll between t = 0 
and t = 24 by the equation: 


f|.z4 + fa-ja (54) 


Solving this differential equation for p and <£, in¬ 
cluding the initial conditions that p = Pi, and 
</> = fa, at t = 0, we obtain: 

V = -^|<i[l-exp(Lp<)-J 

— ( - A[i _ eX p(Lp<)j| + pi exp(L„t) (55) 

0 = + ri 1 - exp( L.o] - i - r p 

- zH 1 ■ expCM ]} 

- P [ 1 - exp(L p <)] + 01 (56) 

At t = 24, let p = P 2 , and <f> = <t> 2 . Then p 2 and 
</> 2 will be given by the following equations: 


KV 1 1 - exp (2L P 4)”1 
th " LX + L p 1 + exp (2L„<i)J 

K[ 1 1 - exp (2^,1 
02 “ LI L 1 L p 1 + exp (2L„<i)J 


(57) 

(58) 


Since we have assumed a steady-state condition, 
the values of p and <f> at t = 24 must be equal in 
magnitude and opposite in sign to their values at 












344 


APPENDIX B 


t = 0. Putting pi = —p 2 and = —<t> 2 , we obtain 
the following equations for p and <f>: 


p= ~rX h ~ t 

_1 1 + exp (2L p ti) — 2 exp ( L p t ) 

L p 1 + exp (2 L p t) 


4> = 


LI 


[ 


1 


H- 1 -^ 


1 1 + exp ( 2L p ti ) — 2 exp(Lp^) 


L x 


1 + exp (2L p h) 


] 


(59) 


(60) 


equations (60) and (61) in equation (64), we obtain: 
Kg 


LphAt — A£ -f- Lp~- -J- 


(AO 2 

2 ' L 


_ K tan a r 

~l7~l 


- i - — h-\-At— -p- 


a 

-rJ-° 


(65) 


Solving for h, we obtain: 

M-J-+ vi~t - ( At ~ ~ 7 ") 

L p T L p tan oc\ 2 L P J 


i + 


9 


VL d tan 


-(1 - L p At) 


- ( 66 ) 


Let us assume that there is a time lag At between 
the time the angular velocity to which the gyro is 
sensitive ( d\J//dt + d<f>/dt tan a) is reduced to zero 
and the time the ailerons reverse their motion. This 
time lag includes the time it takes for the gyro con¬ 
tacts to reverse and the servo clutches to engage and 
reverse the direction of motion of the ailerons. If 
reversal of the aileron motion takes place at t = 2ti, 
the angular velocity to which the gyro is sensitive 
must reduce to zero at time t = 2ti — At. Let the 
values of p and <f> at this time be denoted by p 3 and 
<t> 3 , which are given by the following equations: 

Pz = |-<1 + A< 

_Li 1 - 2 ex p L ,(2h - t) + exp ( 2 L„<i) ~|\ 

L p L 1 + exp (2L P U) .'I l j 

<t>3 — -jrf |b(l — L p t) — At + L p -~- 

_ f" 1—2 exp L p (2ti — t) + exp (2L p ti) 

L P L 1 + exp ( 2 L *i) 

Since, at time t = 2i\ — At, the angular velocity 
to which the gyro is sensitive is reduced to zero, we 
have the following relation: 

^ ^ tan a = 0 at t = 2h — At (63) 



If again we neglect the effects of sideslip velocity, we 
may make use of equation (34) and thus obtain the 
following relation between p 3 and <f> 3 : 

~<t> 3 + (tan a)p 3 = 0 (64) 


Because of the high damping in roll, the terms 
exp (2L p ti) and exp [L p (2h — A2)]are negligible com¬ 
pared with unity for the observed values of h and At, 
and will be neglected in what follows. Substituting 


It is seen from the above equation that for values 
of a such that —g/(VL p tan a) « 1, the period of 
the roll hunt is determined chiefly by the damping 
acceleration in roll L p and the time lag At. The ampli¬ 
tude of the roll hunt from equation (60) is seen to be 
a' linear function of the period and directly propor¬ 
tional to the ratio of the rolling acceleration produced 
per second by the elevons. Thus, to keep the roll 
amplitude small, time lags in the gyro and servo 
units must be kept to a minimum, and the rate of 
application of restoring moment small. 

As a takes on smaller and smaller values so that 
—g/(VL p tan a) becomes of the order of unity, the 
denominator in equation (66) becomes very small, 
increasing the period and amplitude of the roll hunt¬ 
ing motion, until at the value of a determined by the 
equation 

+AI ) = l < 67 > 

the denominator of equation (66) becomes zero, and 
the hunting motion becomes unstable. 

Let us substitute appropriate values of L v , g, V, a, 
and At for SWOD Mark 9 and determine the result¬ 
ant values of the amplitude and period of the roll 
motion. Let us use the values of L p and V given by 
equations (25) and, in addition, set a = 20 degrees 
and At = 0.1 second. Substituting these values into 
equation (66), we obtain h = 0.35 second. Thus the 
roll hunting motion should have a period T = 4/, 
= 1.4 seconds. The amplitude will be approximately 
given by <f> 3 , which has a value of 0.095 radian or 5.4 
degrees. These calculated values check well with ex¬ 
periment as is seen by referring to the curve of Figure 
1 , showing the angle of roll as a function of the time 
for a typical flight of SWOD Mark 9. 

From equation (67), the smallest value of a that 
may be used before the motion becomes unstable is 
about 3.5 degrees. This equation, however, neglects 
rolling moments due to sideslip and to rate of yaw 
which, if taken into account, have the effect of in¬ 
creasing this minimum angle. 
















MODEL TESTS OF LATERAL CONTROL SYSTEM 


345 


B« MODEL TESTS 

OF LATERAL CONTROL SYSTEM 

To study the effect of time lags and speed of the 
servo system on the roll hunting motion, a mechan- 



Figure 3. Mechanical model to represent equation 



Contact spacing-wide, servo speed-normal 



Contoct spocing-moderate, servo speed-normol 


ical model to represent equation (53) was construct¬ 
ed. A photograph of the model is shown in Figure 3. 
It consists of a table free to rotate about a vertical 
axis and carries the turn gyro used in the control 
system. On the same axis is a cylinder which rotates 
inside a concentric cylinder with a small separation. 
The space between the cylinders is filled with oil to 
produce viscous damping of the motion of the table. 
Since the spacing of the cylinders is small and the oil 
is sufficiently viscous, the damping force is quite ac¬ 
curately proportional to the first power of the angular 
velocity of the table. A cord is wound around the 
shaft which supports the table, and each end is con¬ 
nected to a spring. The other ends of the springs are 
connected by cords around pulleys to the servo arms. 
Thus, if the small spring displacement caused by the 
motion of the table is neglected, the torque applied to 
the table will be proportional to the displacement of 
the servo arms and to the constants of the springs. 
If the ratio of the torque applied per radian displace¬ 
ment of the servo to the moment of inertia of the 
table is made equal to the value of Lj for the missile, 
and the ratio of the damping torque to the moment of 
inertia of the table is made equal to the value of L p 



Contact spocing-close, servo speed-0.7 X normol 



Contoct spocing - close, servo speed- normol 



Contoct spacing-close, servo speed-normol Contoct spocing-close, servo soeed - 1.3X normol 

Figure 4. Roll simulator records showing effects of gyro contact spacing and servomotor speed on amplitude of roll 
hunting motion. Sawtoothed type curves represent differential motion of elevons; sinusoidal type curves, angular motion 
of table; and broken lines, operation of right and left servo clutches, respectively. Film speed is approximately 1 inch per 
second. 












346 


APPENDIX B 


for the missile, the motion of the table will be gov¬ 
erned by equation (53). If the turn gyro mounted on 
the table is connected to the servo, as in the missile, a 
hunting motion will be set up, simulating the roll 
hunting motion of the glider. Some records obtained 
with the model are shown in Figure 4; they were 
taken to study the effect of time lag in the system and 
rate of movement of ailerons on the amplitude of the 



Figure 5. Series of runs in which angle of gyro is 
varied from 2.5 to 20 degrees. 


roll hunt. The results are in accord with the theoreti¬ 
cal treatment above. 

The Servomechanisms Laboratory of the Massa¬ 
chusetts Institute of Technology, in connection with 
the development of an alternative control system for 
SWOD Mark 7 and Mark 9, constructed a device to 
simulate the roll and yaw motions of a glider. A plat¬ 
form was arranged so as to be free to rotate about a 
horizontal axis (representing roll motion) and to be 
driven in rotation about a vertical axis. A torque 
motor and generator were connected to the roll axis, 
and synchros were attached to the arms of the servo 


under test. Electrical circuits were arranged so that a 
torque was applied about the roll axis proportional to 
the differential displacement of the ailerons and to 
the rate of roll, with proper constants of proportion¬ 
ality so that the motion in roll would be represented 
by equation (53). The platform was driven in rota¬ 
tion about a vertical axis at an angular velocity pro¬ 
portional to the angular displacement about the hori¬ 
zontal or roll axis, and the constant adjusted so that 



its motion would be represented by equation (34). 
All effects due to sideslip and the cross derivatives L r 
and N p were neglected. 

The turn and bias gyros and antenna system used 
in SWOD Mark 7 were mounted on the platform and 
a beacon for homing signals placed at a distance from 
the platform. The system was operated so that the 
lateral motions of the glider in flight would be simu¬ 
lated. Records obtained of the angle of bank <f> and 
the angle of yaw ^ of the platform as a function of the 
time for various adjustments of the control system 
are shown in Figures 5, 6, and 7. In all records 



















































MODEL TESTS OF LATERAL CONTROL SYSTEM 


347 


shown, the platform was initially set 5 degrees off the 
axis of homing. The vertical scale for </> is four times 
that for \J/. 

Figure 5 shows a series of runs in which the angle 
of the gyro is varied from 2.5 to 20 degrees. For 2.5 
degrees the motion is unstable, as predicted by the 
roll stabilization theory in Section B.5, the platform 
hitting its limit stops during the test. The value of 15 
degrees, found to be most satisfactory from flight 
tests, is somewhat larger than the best result from 
the model tests, the difference being due to the terms 
neglected in the equation governing the model tests. 

Figure 6 shows a series of runs in which the angle 
designated as ^ 0 , the minimum error angle that pro¬ 
duces saturation of the differential amplifier which 
feeds into the turn gyro coils, is varied. This, in 
effect, is the same as varying the value of c, the rate 
of yaw in degrees per second that produces the same 
torque on the gimbal frame of the gyro as an angular 
error of 1 degree. The value i/'o = 6 degrees was 
found most suitable from flight tests, and this value 
is seen to be satisfactory by this model test. 

Figure 7 shows the effect of the bias gyro on the 
damping of the yaw oscillations. If a bias gyro is not 
used (R = oo), sustained oscillations in yaw are ob¬ 
tained both in flight tests and in the model test. The 
most suitable value for the biasing resistor deter¬ 
mined from both flight tests and model tests was 
found to be about 5,000 ohms. 
























Appendix C 

INFRARED RADIATION 

AND ITS APPLICATION TO HOMING MISSILES 


Cl INTRODUCTION 

Target-seeking devices based upon infrared radi¬ 
ation are rendered unique by two factors. First, an 
object at any temperature different from its sur¬ 
roundings radiates heat differentially. Second, this 
differential radiation can be detected in the dark 
without revealing the observer, as occurs, for exam¬ 
ple, when using radar. The term “infrared” has been 
used to include all electromagnetic waves between a 
wavelength of 7,000 A (= 0.7ju = 7 X 10~ 5 cm) and 
3,000,000 A (= 300 ju = 0.03 cm). Wavelengths 
greater than this are most satisfactorily produced by 
electrical means. 

In accordance with well-known laws, the total 
amount of energy radiated by any object varies with 
the fourth power of its absolute temperature, and the 
distribution of the radiated energy among the differ¬ 


ent frequencies changes with altered temperatures. 
In Figure 1 is plotted the radiant energy as a function 
of its wavelength for several different temperatures. 
This shows graphically the reason for our particular 
interest in using the far infrared for heat detection. 
It will be seen that the temperature must rise to 
about 500 C (932 F) before there is any appreciable 
radiation appearing at wavelengths visible to the 
eye. Nevertheless, at any temperature above abso¬ 
lute zero, any object is continuously radiating heat 
energy. Furthermore, for many military targets 
whose temperatures will be slightly above that of 
their surroundings, the peaks of the radiation curves 
occur within the wavelength band designated in Fig¬ 
ure 1 as the “water vapor window.” The atmosphere 
is transparent to radiation of these wavelengths, 
while on either side are large areas within which ab¬ 
sorption by atmospheric carbon dioxide and water 



WAVELENGTH IN MICRONS 

Figure 1 . Black-body radiation. 


348 














GENERAL CONSIDERATIONS 


349 


vapor is complete. This transmitted radiant energy 
may be detected by appropriate devices other than 
the eye. The fact that such devices can be made of 
phenomenal sensitivity has led both scientists and 
gadgeteers to pursue intensive work in this field. 
This in turn has resulted in many hundreds of pro¬ 
posed, and a few successful, devices. 

Devices for detecting an object by infrared radi¬ 
ation may be sharply classified into two groups: 
those utilizing the far and those dependent upon the 
near infrared. Partly by nature and partly by defini¬ 
tion, the two classes have the following character¬ 
istics: 


Far Infrared 

1. “Far infrared,” as used 
in this report refers to heat ra¬ 
diations of wavelengths longer 
than 2.5 n . However, because 
of atmospheric absorption, 
operations are restricted to 
the band between 8.5 and 
13 n. 

2. Far infrared devices usu¬ 
ally depend upon the heat 
energy radiated by all objects 
rather than upon the use of 
special sources. Reflection is 
of minor importance. Most 
solids and liquids radiate ap¬ 
proximately as “black bodies” 
at these wavelengths, so that 
the contrast for any detection 
problem is primarily deter¬ 
mined by temperature dif¬ 
ferences. 

3. Most solids and liquids 
absorb completely all radia¬ 
tion not reflected at their sur¬ 
faces, so that very few ma¬ 
terials may be used as lenses 
and windows. Changes in di¬ 
rection must usually be ac¬ 
complished by reflection. 

4. Photoelectric and photo¬ 
chemical receivers cannot be 
used. The basis of operation 
depends on the absorption of 
radiant energy as heat. 


Near Infrared 

1. “Near infrared,” as used 
in this report, refers to the 
wavelength band between the 
visible and 2.5 n. 


2. Most near infrared de¬ 
vices require the use of a spe¬ 
cial source of radiation di¬ 
rected over a path of some 
kind to the receiver. For de¬ 
tection problems, reflected 
radiation is frequently used. 
With only a few exceptions, 
when a natural source is hot 
enough to emit much energy 
in the near infrared, it is hot 
enough to emit some in the 
visible at the same time. 

3. The optical system may 
be handled by the same gen¬ 
eral methods as for visible 
light. Certain selective absorp¬ 
tions and reflections create 
some of the special uses for 
near infrared devices. 

4. The most sensitive and 
rapid receivers in this range 
are photoelectric and photo¬ 
chemical devices. 


C 2 GENERAL CONSIDERATIONS 

Many of the problems met in infrared research are, 
of course, common to both the above categories. In 
this classification falls the matter of attenuation over 
long paths in the atmosphere. The principal constitu¬ 
ents of the atmosphere, nitrogen and oxygen, have no 
absorption bands of any importance to this discus¬ 
sion. On the other hand, water vapor, carbon dioxide, 


and ozone have strong absorption bands, and various 
atmospheric colloidal particles—smoke, dust, clouds, 
etc.—cause considerable attenuation. 

c 21 Energy Loss 

Due to Tiny Suspended Particles 

Any solid or liquid particles in the atmosphere may 
attenuate incident radiation by two distinct proc¬ 
esses. The first is by absorption of the radiant energy 
within the particle, which in general will result in its 
reradiation at a longer wavelength corresponding to 
the black-body temperature of the particles. The sec¬ 
ond is by a change in the direction of the incident 
radiation due to diffraction, reflection, and refrac¬ 
tion, or to any one of these. Such change in direction 
after contact with many particles is considerable, and 
virtually no radiation is transmitted. Only when the 
particle size is small with respect to wavelength, as is 
the case in very faint haze, do the dispersing phenom¬ 
ena become substantially less effective. When the 
particle size is very small, attenuation may be due 
only to scattering. In such cases, as is well known, the 
scattering may be considerably less for the longer 
waves. 

There are essentially two types of atmospheric 
particles—solids and liquids. 

Smoke, Dust, etc. Solid particles are nearly opaque 
to visible and infrared radiation; hence, attenuation 
by large particles is due simply to reflection and ab¬ 
sorption. For particles small with respect to wave¬ 
length, attenuation is chiefly by scattering. The re¬ 
sulting use of near infrared in aerial photography for 
penetration of atmospheric haze is common knowl¬ 
edge. 

The penetration of smoke screens is another appli¬ 
cation for infrared which has received some consider¬ 
ation. A limited number of field and laboratory tests 
on smoke penetration have been made. 1 These indi¬ 
cate that there is good penetration by the far infrared 
in some cases and by the near in a very few cases, 
depending on the type of smoke used. In general, the 
size of the smoke particles tends to increase with 
time after original production of the smoke, so that 
penetration, particularly in the near infrared, de¬ 
creases eventually to useless values. Attenuation by 
other opaque particles, such as dust, is also a matter 
of particle size and presents a similar special problem 
in each case considered. 

Water Droplets. A considerable study, by both ex¬ 
perimentation, 1,2 and theory 3 has been made of infra¬ 
red fog and cloud penetration. The evidence at first 
seems conflicting. However, if the question of particle 
size is correctly understood, the results are conclu¬ 
sive. This quantity is difficult to measure, particular¬ 
ly when the density of a fog or haze is so low that the 



350 


APPENDIX C 


visibility may be several miles. However, transmis¬ 
sion for several wavelength bands and the distribu¬ 
tion of water particle sizes have been measured simul¬ 
taneously in a series of experiments. 4 The longest 
wavelength band used included the 8.5- to 13 -m water 
vapor window. Something of the order of one hun¬ 
dred observations were made in several different nat¬ 
ural fogs and clouds. In no instance was any differ¬ 
ence in transmission observed between the visible 
and the infrared. All this information indicates con¬ 
clusively that reflection, refraction, absorption, etc., 
are sufficient at all useful wavelengths to restrict the 
transmission in most fogs and clouds to ranges little, 
if any, greater than for visible light. Use of the infra¬ 
red beyond 15 y is not helpful, as water vapor ab¬ 
sorbs strongly at longer wavelengths, and energy 
levels are low. 

When the particle size is small with respect to 
wavelength, the problem is the same as for dust, 
smoke, or any other small particles. Apparently, haze 
and fog particles often are less than 1 y in size when 
the visibilities are a mile or more. In such a haze 
infrared might have appreciably increased range over 
the visible. Some dense fogs and clouds may have a 
sufficiently high percentage of small droplets so that 
a small increase in range for the infrared might be 
observable through them, also, and some fogs may 
contain a large amount of smoke. However, it is un¬ 
usual to find a cloud in which more than 50 per cent 
of the particles are smaller than 10 y in diameter or a 
sea fog in which many are smaller than 20 y. A In most 
clouds and fogs the infrared is attenuated as much as 
the visible. 

Mist and raindrops (also snowflakes) are, of course, 
considerably larger than any of the above, and no 
reliable observations in the infrared indicate any im¬ 
provement over visible light when such large parti¬ 
cles only are in the optical path. 

Certain experimental results obtained under spe¬ 
cial conditions, and some theoretical work based on 
rather doubtful assumptions, indicate some improve¬ 
ment in transmission in the infrared. This work 
accounts for the common misunderstanding concern¬ 
ing fog and cloud penetration. It is considered, how¬ 
ever, that if the visual range is something of the 
order of 500 yd, the complexity of infrared apparatus 
needed to increase that range to 600 or even to 800 yd 
is not justified. 5 


022 Absorption Bands Due to Gases 

Water Vapor. The absorption by water vapor has 
been measured in detail throughout the infrared 
spectrum over actual atmospheric paths. 6 Under av¬ 
erage conditions, the total amount of water vapor, in 


a path a few hundred yards long near the surface of 
the earth, is large enough so that absorption may be 
considered complete at certain wavelengths for such or 
longer paths. 

Carbon Dioxide. This gas is present to an extent of 
but 0.03 to 0.04 per cent, but this is sufficient to cause 
sharp absorption at several wavelengths in the infra¬ 
red. 7 

Ozone. Substantial quantities of ozone in the strato¬ 
sphere limit the infrared and the ultraviolet solar 
radiation received at the earth’s surface. Below 
20,000 ft, however, ozone is negligible. 

Absorption by water vapor and carbon dioxide in 
the atmosphere generally limits the use of infrared to 
two broad bands in the spectrum, one extending from 
the visible to about 4 y and the other from 8.5 to 
13 y. The former covers the near infrared, plus a por¬ 
tion of the far infrared which does not seem particu¬ 
larly useful for military purposes. The temperatures of 
sources whose radiation peaks are between 2.5 and 
4 y (see Figure 1) range from about 500 C to 1,000 C. 
Such sources seem to be uncommon, and where they 
do occur they are easily shielded. Furthermore, those 
in the higher part of this range radiate sufficiently in 
the near infrared to be detected by the more sensitive 
photoelectric receivers. As is indicated in Figure 1, 
increasing the temperature of a black body increases 
the intensity of radiated energy at all wavelengths at 
which there is any emission, not just at the wave¬ 
length of the peak. 

Water vapor shows strong absorption between 13 
and 20 y, and beyond 20 y it renders the atmosphere 
virtually opaque to infrared radiation. In general, 
therefore, the nature of the sources, existing receiv¬ 
ers, and atmospheric absorption confine the work to 
the near infrared (0.7 to 2.5 y) or to the 8.5- to 13 -m 
water vapor window. 

As will be seen from Figure 2, this limitation is not 
particularly unfortunate. Here is shown the per cent 
transmission at various wavelengths through an at¬ 
mospheric path containing water vapor and carbon 
dioxide. On the same wavelength scale are super¬ 
posed the curves for black-body radiation at temper¬ 
atures from 0 C to 100 C. It is evident that this win¬ 
dow covers the peaks of the radiation curves for ob¬ 
jects at common terrestrial and atmospheric temper¬ 
atures. In this region, also, nonselective receivers 
may still be very sensitive. 

For aerial photography, where atmospheric haze 
presents a common difficulty, near infrared suffers 
less attenuation than visible light. Other than this, 
the use of near infrared offers transmission advan¬ 
tages over the visible only in very special cases. 
Smoke-screen penetration might be an example. The 
8.5- to 13-/x transparent band, however, is useful be¬ 
cause it covers exactly the wavelengths radiated by 



HEAT-DETECTION DEVICES 


351 


many desirable military and naval targets under 
usual conditions. 

0,2-3 Emission of Infrared Radiation 

The most common sources of infrared radiation are 
solid materials which, in general, radiate as black 
bodies. In other words, their surfaces may be consid¬ 
ered 100 per cent absorbing or 100 per cent emissive 
at any infrared wavelength. Most solids and liquids 
come closest to being ideal black bodies at the longer 
wavelengths, beyond 5 or 10 /jl. Exceptions are pol¬ 
ished metals, such as are used for reflectors. The most 
common sources in the near infrared are simply solid 
materials heated to incandescence. 

Gases behave in a very different manner from sol¬ 
ids and liquids, radiating and absorbing only in 
rather narrow spectral bands. It has frequently been 
suggested that various targets, particularly aircraft, 
be detected by the energy radiated by their exhaust 
gases. However, exhaust gases are largely water va¬ 
por and carbon dioxide, which are also common in 
the atmosphere. Consequently, the energy radiated 
at wavelengths characteristic of these gases is re¬ 
absorbed in a relatively short atmospheric path. This 
conclusion has been verified experimentally. 

Where long-wavelength infrared is concerned, it is 
important to remember that almost all solid and liq¬ 
uid objects are radiating in a manner characteristic of 
their temperature. Consequently, any infrared re¬ 
ceiver is affected by radiation from its immediate 
surroundings and from any other object included in 
its field of view, as well as from the particular source 
under consideration. Any heat source must be con¬ 
sidered with relation to its background. This point is 
fundamental in designing infrared receivers. 

0-2,4 Conclusion 

Any solid body which is at a temperature higher 
than its surroundings (and this includes many other¬ 
wise indistinguishable military and naval targets) 
will emit heat radiation in the far infrared which can 
be detected by appropriate receivers at distances of 
several miles through clear or hazy air. 


c 3 HEAT-DETECTION DEVICES 

Reasons have already been advanced for the use of 
the far infrared between 8 and 13 /x (80,000 and 
130,000 A, 0.0008 and 0.0013 cm) for detection of 
military targets. It has been shown that such varied 
targets as men, ships, buildings, and airfield landing 
ramps are continuously radiating energy which can 
be detected at distances up to several miles by the 


use of appropriate devices. The chief limitation in 
detection is the factor of background radiation, an 
object being detectable only if it stands out suffi¬ 
ciently against a more or less uniform heat back¬ 
ground. This memorandum describes means for de¬ 
tecting such differential radiation. 

0,3,1 Receivers 

The utility of any device is obviously determined 
by the characteristics of the receiver. Any far infra¬ 
red receiver is essentially a radiation thermometer, in 
which radiant energy must be absorbed by the sensi¬ 
tive element, producing a temperature change. Meas¬ 
ures of merit which should be considered in compar¬ 
ing different receivers are: ( 1 ) sensitivity; ( 2 ) speed 



Figure 2. Black-body radiation and relative trans¬ 
mission through air containing H 2 0 and C0 2 . 


of response; (3) ruggedness; (4) adaptability to a 
suitable type of indicator; (5) size and shape of sensi¬ 
tive area; ( 6 ) freedom from vibration, from changes 
in ambient temperature and pressure, etc.; (7) size 
and weight, simplicity of essential auxiliary appara¬ 
tus, etc.; ( 8 ) other special features. 

Determination of relative merits of receivers, 
therefore, is truly valid only in terms of specific ap¬ 
plications, but it is possible to make a few general 
comparisons. 

The temperature change produced by the ab¬ 
sorbed radiation may be measured conveniently by 
any of three well-known types of sensitive receiver. 

Thermocouples. A thermocouple is a joined pair of 
conductors of dissimilar metals forming a closed cir¬ 
cuit. Any difference in temperature between the two 









352 


APPENDIX C 


junctions will produce a measurable voltage differ¬ 
ence, and hence, a current will flow through the cir¬ 
cuit. Several tiny couples may be connected in series 
to increase the sensitivity, forming a thermopile. 

Thermocouples may be made of fine wire or of thin 
films of metal on a nonconducting surface, such as 
cellophane. The speed of response is governed by the 
mass of material to be heated. In most cases with 
which we are concerned, this means that the thermal 
mass should be as small as possible. Thermocouples 
may be made of approximately the same sensitivity 
and ruggedness as any other device but tend to be 
slower in response. Suitable amplification for indica¬ 
tion can be provided. Size and shape are less variable 
than is the case for bolometers and gas radiometers. 8 

Bolometers. A bolometer is a radiation receiver 
whose temperature change due to incident radiation 
is observed through some accompanying change in 
its electrical properties, such as resistance or dielec¬ 
tric constant. 

Here again, the smaller the thermal mass, the more 
sensitive the device and the more rapid its response. 9 
Bolometers have been made of thin films or strips of 
metal, of semiconductors with high temperature co¬ 
efficients, 10 or of condensers whose dielectric constant 
changes with temperature. 11 

Bolometers of special construction have been made 
with extremely high sensitivity, high speed of re¬ 
sponse, adequate ruggedness, and freedom from un¬ 
wanted disturbances. The signal from bolometers is 
particularly well adapted to electronic amplification. 

Gas Radiometers. If radiation is admitted to a tight, 
gas-filled chamber in which the radiation is absorbed, 
the temperature, and consequently the pressure, will 
be raised. The pressure change may be observed op¬ 
tically or electrically. 12 

Such instruments may be rugged, extremely sensi¬ 
tive, and rapid in response, producing a signal capa¬ 
ble of high amplification, but they are subject to 
greater difficulties from vibration than the other 
instruments. 

Image-Producing Receivers. So far, the only method 
of producing an image or “infrared picture’ ’ has been 
by some scanning process, except for some instru¬ 
ments which are as yet in the early research stage. 13 


c * 3,2 Applications of the Far Infrared 

This section has been interested in the application 
of the far infrared to automatic control of homing 
missiles. Other workers have investigated possible 
uses in signaling and in various detection problems. 
Under proper conditions the latter idea shows 
promise. 

Automatic Control. By the use of electronic amplifi¬ 


cation, the signals from infrared receivers can be used 
for relay operation and consequently can serve to aim 
or guide a vehicle toward any heat source or other 
thermal discontinuity in the background. 

This section has concentrated its efforts in the heat 
field on the development of a high-angle, heat¬ 
homing bomb, designated “Felix.” 

This consists of a standard 1,000-lb GP demolition 
bomb with added false nose and tail. The nose carries 
the heat-detection unit with its associated amplifier, 
while the tail carries the means of guiding the flight. 
A small, parabolic mirror mounted with its axis 5 de¬ 
grees off center is rotated rapidly, thereby scanning a 
field totaling 20 degrees in diameter. Any thermal 
discontinuity in the field produces an electrical pulse 
in the output of the nickel-strip bolometer. A com¬ 
mutator on the mirror drive shaft makes possible 
identification of the particular quadrant of the field 
in which the target is located. Suitable relays and 
servomotors operate the control surfaces of the bomb 
in accordance with the information received from the 
eye. As an antihunt device, the whole eye is linked 
with the rudder in such a manner that the projectile, 
in turning, is made to see ahead of where it is going at 
the moment. Only when it is pursuing a straight path 
does it look directly forward. A bolometer strip of 
blackened nickel, 0.2 n to 3 n thick, 0.25 mm wide, 
and 5 mm long, in low-pressure hydrogen, gives 
adequately fast response. 


033 Background Radiation 

It is obvious that any automatic control problem 
involves discrimination between the particular target 
and its surroundings or background. If variations in 
the background are of the same order of magnitude 
as the discontinuity presented by the target, such 
means of control cannot be used. It must be remem¬ 
bered that any single receiver has a field of view and 
that the field of view of such a single element deter¬ 
mines the resolution of the whole device. Conse¬ 
quently, even if a particular target is very much hot¬ 
ter or colder than any other area in the background, 
if it is small with respect to the optical resolution of 
the receiving system, it may be impossible to dis¬ 
tinguish it from the background. In other words, the 
change in radiant energy at the receiver as it sweeps 
across the target may still be less than changes pro¬ 
duced by the background. Using a smaller angle of 
view lessens this particular difficulty but at the same 
time increases the difficulty of searching. A number 
of receivers are now capable of detecting the radi¬ 
ation from common military targets at distances up 
to several miles. When these detectors are inserted in 
heat-homing missiles, the necessarily widened field of 



HEAT-DETECTION DEVICES 


353 


view makes the target-to-background signal ratio or 
signal-to-noise ratio, rather than ultimate sensitivity, 
the limiting factor. 

Studies on background radiation have been carried 
out extensively by Section 16.4 of NDRC and by the 
U. S. Navy Department. In general, natural vegeta¬ 
tion tends to remain at the same temperature as the 
surrounding air, both day and night. Consequently, 
regions well covered with grass and trees present a 
fairly uniform background. By contrast, in urban 
areas surface temperatures vary widely: in the day¬ 
time, because of varying rates of absorption of solar 
energy; at night, because of varying rates of cooling 
and different internal heating. 

The conclusion is that satisfactory heat targets are 
moderately rare on land. Large factory buildings or 
areas are acceptable when surrounded by natural 
vegetation, but usually not when in the midst of 
urban districts. The influence of camouflage on this 
picture might be interesting. Airport landing strips 
and parking ramps offer one particularly useful field 
for attack by heat-homing missiles. On the other 
hand, ships of 5,000 tons and up may be detected 
easily from altitudes of 10,000 ft against sea back¬ 
grounds. 


c 3 4 Methods of Operating 

Infrared Receivers 

This section will compare two rather different 
methods of operating long-wavelength infrared re¬ 
ceivers for detection or control. It is obvious from the 
preceding section that any detection or control appa¬ 
ratus functions best on the differential between the 
target and its background, not on the absolute signal 
strength of the target. This may be accomplished in 
two ways: la 

1. Steady-state method. Provide two receivers. One 
“sees” the target, the other its adjacent background, 
and the indicator operates on the difference in tem¬ 
perature between the two receivers. 

2 . Pulse method. Provide one receiver, which sees 
the target and its background alternately, and an 
indicator operating on the resulting change in tem¬ 
perature of the receiver. 

Consider the use of radiation thermopiles for spe¬ 
cific examples of each method. In the first, two re¬ 
ceivers would be provided and the thermocouple 
junction connected so that the emf would be in one 
direction if one receiver was warmer than the other, 
and the reverse if colder. The two receivers would be 
mounted adjacent to each other in the focal plane of 
a mirror or lens so that an image of the target could 
be placed on one receiver and an image of the back¬ 
ground on the other. Any resulting emf would then 


be amplified by some means and used to aim a de¬ 
tector or operate a control. 

In the second method, a single receiver would be 
placed in the focal plane of the optical system and 
one set of thermopile junctions connected to this re¬ 
ceiver, the other set to a large thermal mass at am¬ 
bient temperature. In such a system, under most 
conditions, there would be a difference between re¬ 
ceiver and ambient temperature, and a resulting emf, 
but the coupling to the amplifier should be made so 
that only changes in emf would cause signals. Some 
sort of scanning would be provided to produce a 
change in emf when the image of the target crossed 
the receiver. 

Other types of radiation receivers may also be used 
in both ways, and thermopiles may perhaps be used 
more effectively in some variation of the above, but 
here we are comparing only the two methods. The 
first permits radiation from the target to remain on 
the receiver for a comparatively long period of time 
so that, in general, radiation equilibrium may be 
established with the target; this is known as the 
steady-state method. The second, involving scan¬ 
ning, places only a pulse of radiant energy on the 
receiver and has been named the pulse method. 

The chief advantage of the steady-state method is 
that it permits the realization of higher sensitivity 
simply because the longer exposure times permit the 
receiver to absorb more energy. The chief advantage 
of the pulse method is that any false signals super¬ 
imposed on the system may be excluded from the 
amplifier, provided they have a long period compared 
to the rate of scanning. The most sensitive steady- 
state receivers devised to date have a relaxation time 
of 0.5 to 10 seconds; i.e., that amount of exposure 
time is required to approach the steady-state condi¬ 
tion. Many military detection and control problems, 
however, require that an exposure time of consider¬ 
ably less than 1 second produce usable signals. Also, 
sensitivity is greatest for an optical system which 
just covers the receiver with an image of the target; 
but again, speed of operation requires a considerably 
greater field of view. These two factors make it diffi¬ 
cult for the steady-state method to be fully effective. 

The pulse method, on the contrary, must be based 
on the use of a fast-response receiver. Where a relax¬ 
ation time t of 1 second might be satisfactory for a 
steady-state receiver, from 0.1 to 0.001 second would 
be required for one working on the pulse method. 
Sensitivity is proportional to y/t, other factors being 
properly matched. The steady-state method would, 
then, be three to thirty times more sensitive than the 
pulse method, because of this effect. However, the 
scanning procedure used in the pulse method permits 
a smaller instantaneous field of view, which gives a 
comparative gain in sensitivity if the total field of 




354 


APPENDIX C 


view swept out by the scanning equals the stationary 
field of the steady-state receiver. The ratio between 
the two instantaneous fields might also be three to 
thirty times, so that these two factors offset each 
other. It is concluded, then, that the maximum sensi¬ 
tivities which can be achieved by the two methods 
are not widely different for most military applica¬ 
tions. There will, of course, be exceptions. 

The greatest advantage of the pulse system has not 
yet been considered; namely, its elimination of slow 
“drift” effects. All slow changes in the receiving net¬ 
work of a steady-state system appear as signals. In 
the pulse system, however, only changes which occur 
at the scanning frequency come through as signals. 
The causes of false signals which may occur in the 
two systems will now be considered. 

In steady-state systems two receivers and their 
circuits are compared, so that any difference along 
the network causes an apparent change in the match¬ 
ing. Such changes might be caused by thermal emf’s 
at points other than at the thermocouple junctions, 
or resistance changes if we have a bridge circuit, or 
other effects of changing temperature and pressure, 
which will vary in each case with the particular type 
of receiver and circuit. Also, if the radiation receivers 
are not perfectly matched with respect to the ther¬ 
mal mass, their temperatures will change at different 
rates as the ambient temperature changes or as they 
may shift from a condition of radiation balance with 
a very cold field of view to radiation balance with a 
field considerably warmer. 

To overcome these difficulties, the circuit must be 
very carefully designed and matched electrically, the 
entire circuit well insulated thermally and possibly 
held at constant temperature by a thermostat. Even 
so, all steady-state systems must be furnished with 
some balancing adjustment. This adjustment can be 
either manual or automatic; if automatic, it can be 
arranged to change the balance at such a rate as to 
overcome changes which occur at a sufficiently slow 
rate. The system then begins to approach the pulse 
method, but in a rather complex manner. 

In the pulse method, slow changes have no effect 
unless they are so large as to affect the sensitivity by 
severely unbalancing a bridge circuit, overloading a 
transformer with direct current, blocking the ampli¬ 
fier, or some similar effect. There is, however, a diffi¬ 
culty to which the steady-state system is not subject 
at all: the circuit is commonly tuned to the scanning 
frequency, or at least responds to signals occurring at 
that frequency, and in that case some false signals 
are bound to be generated by the scanning itself. For 
example, microphonics resulting from mechanical un¬ 
balance might be synchronous with the scanning fre¬ 
quency. Various other kinds of noise due to the scan¬ 
ning may occur. 


It is essential to consider all the facts mentioned 
above in designing an infrared detector or control 
apparatus. In general, for an application which per¬ 
mits fairly slow operation, such as searching for ships 
from a shore station, the steady-state method may be 
used if the apparatus can be kept in balance. For 
applications involving very rapid response, as is the 
case with most airborne equipment, the pulse method 
appears more promising. 

0,3 5 Infrared Receiver Development 

This report covers the receiver development which 
has taken place under Section 5.5 of NDRC toward 
two objectives. The first was to develop a receiver 
particularly adaptable to flight conditions, including 
flight in a target-seeking bomb. This requires extreme 
portability, rapid response, large angle of view, free¬ 
dom from microphonics and from drift of “zero set¬ 
ting,” and operation throughout rapid changes in 
ambient temperature and pressure. It was felt that 
these requirements could best be met by a simple 
metallic bolometer in a d-c bridge circuit with some 
sort of scanning to furnish a-c pulses for electronic 
amplification. Work has also been done with thin 
thermopiles produced by evaporation. 

The second objective was to develop receivers with 
similar characteristics but with considerably higher 
sensitivity than anything previously available. A 
very promising possibility was found in the ther¬ 
mistor bolometer initiated by the Bell Telephone 
Laboratories and further developed under the super¬ 
vision of Section 16.4 of NDRC. 

Metallic Bolometers 

The Hammond Platinum Bolometer. Starting in 
January 1941, Laurens Hammond, of Chicago, co¬ 
operated in the development of an infrared device for 
locating surface vessels from aircraft. As primary re¬ 
ceiver, Hammond used a thin platinum-strip bolom¬ 
eter. When two platinum strips about 1 ju thick, 
connected in adjacent arms of an a-c bridge, were 
used, moderate sensitivity was obtained. Later in 
1941, it was decided that the use of a d-c bridge cir¬ 
cuit was much more satisfactory for an instrument 
requiring extreme portability. When the circular 
scanning system already mentioned was used, any 
thermal discontinuity in the field created an elec¬ 
trical pulse in the bridge circuit. By this means an a-c 
amplifier could be used with the d-c bridge circuit. 

As this work progressed, it became apparent that 
both higher sensitivity and faster response would be 
essential. 

Improved Metallic Bolometers. Under Contract 
OEMsr-60, Harvard University 10 has been able to 



HEAT-DETECTION DEVICES 


355 


improve greatly the design of metallic bolometers 
requiring rapid response. It was shown that a small 
thermal mass is essential to good sensitivity. Other 
considerations in the same study indicate that the 
shorter the relaxation time of the bolometer, the 
greater the sensitivity. Designed on the basis of the 
factors discussed above, bolometers have been made 
with relaxation times as short as 0.002 second. Sub¬ 
stitution of the thin nickel strips for platinum, and 
mounting them in an atmosphere of hydrogen under 
a pressure of 5 mm has resulted in bolometers with 
sensitivity enough to detect a minimum signal of 
10 -8 watt per sq cm when used with the scanning 
system employed for Felix. 

Much acrimonious dispute has raged over the rel¬ 
ative utility of these different types of receiver. As 
has been pointed out, the absolute sensitivity is not 
the only criterion. If there were no other limitations, 
we could amplify the receiver voltage indefinitely, 
and hence, we could detect indefinitely small quanti¬ 
ties of heat. However, it is well known that a certain 
random noise voltage will be developed across the 
grid of the first tube in the amplifier. If the voltage 
across the bolometer strip is much less than this, it 
will be swamped by this noise. Noise voltage may be 
due to (1) fluctuations arising in the amplifier, 
(2) thermal noise in the bolometer bridge—so-called 
“Johnson noise,” (3) “current” or contact noise and 
(4) microphonics, if any. By proper design it is pos¬ 
sible to minimize, but never to eliminate, this irregu¬ 
lar voltage variation. In addition, of course, for 
homing devices the thermal noise due to response to 
the background is high. Subsequent amplification 
applies both to target-signal voltage and to noise 
voltage, so that the limiting factor in determining 
useful sensitivity is the ratio of signal voltage to 
noise voltage. For control purposes, this ratio must 
be of the order of three to one or larger, or erratic 
operation may ensue. 


036 Infrared Optics 

Scanning Systems. For automatic control of a mis¬ 
sile, a circular scanning pattern seems necessary, 
rather than the simplest form of back-and-forth line 
scanning. Most of the work of this group has been 
carried out with a system which scans by rotating a 
parabolic mirror off axis. The alignment is such that 
the axis of rotation passes through the optical axis at 
the focal point of the parabola. This means that an 
image at the center of the receiver is always in sharp 
focus. Such a system can be arranged to scan a circu¬ 
lar area and, by means of a commutator, to indicate 
in what sector of the field of view any particular heat 
source lies. 


Transparent Materials. The classic material for far 
infrared optical parts has been rock salt. The only 
limitations in its use are the difficulty of obtaining it 
in large pieces, and its solubility in water. Waterproof 
coatings have been necessary to combat the latter. 
For Felix, however, it has been found that sheet sil¬ 
ver chloride possesses many advantages. Experi¬ 
ments indicate that silver chloride sheets can be cut 
and pressed into the desired shapes for windows 
which are thin enough so that rapid changes in am¬ 
bient temperature do not produce fogging, strong 
enough to resist wind velocity of the magnitude to be 
expected, and yet transparent enough to transmit at 
least 80 per cent of the incident energy in the 8.5- to 
13 -m band. 

0,3,7 Infrared Electronics 

Almost all infrared devices considered by this sec¬ 
tion have involved electronic amplification. As has 
been pointed out, most of the work has been with 
resistance bolometers placed in some sort of bridge 
circuit. This leads immediately to two quite different 
means of amplification, both of which have been ex¬ 
tensively developed. The first involves the use of an 
a-c bridge current and an amplifier tuned to the same 
frequency. The second uses a d-c bridge current with 
an amplifier less sharply tuned to amplify pulses of 
unbalance of the bridge. 

A-C Bridge System. This system has been used al¬ 
most entirely for the thermistor bolometer and is 
under development. 



D-C Bridge System. A major difficulty in the design 
of an amplifier for the d-c bolometer excitation sys¬ 
tem is the building of a low-frequency amplifier 
which will give a sufficiently good ratio of signal to 
noise and which at the same time does not produce 
too much phase shift with variations in scanning fre¬ 
quency. It may be impossible to obtain as high over¬ 
all sensitivity by this method as is possible with an 
a-c bridge system, but there are other advantages for 
a rugged field instrument. These advantages have 
been brought out. A large part of this work has been 
done in designing such an amplifier for use with the 
















356 


APPENDIX C 



Figure 4. Wiring diagram of electronic unit. 































































































































































































































CONCLUSIONS 


357 


low-resistance metallic bolometer. Because of its low 
resistance, the bolometer is coupled to the amplifier 
by means of an input transformer. The most efficient 
input circuit seems to be that shown in Figure 3, 
which is the result of a careful theoretical analysis of 
the network. The amplifier has been particularly de¬ 
signed for control relay operation so that it is quite 
different from one which might be evolved for a dif¬ 
ferent purpose. (See Figure 4.) 

c.4 CONCLUSIONS 

In brief, then, Section 5.5 of NDRC has demon¬ 
strated that many military and naval targets are con¬ 
tinuously radiating large quantities of energy within 


the wavelength band from 8.5 to 13 /j, to which the 
lower atmosphere is transparent. These wavelengths, 
far beyond the visible, will pass through most ordi¬ 
nary military smoke screens, and, of course, are as 
easily received by night as by day. The use of easily 
produced metallic bolometers has demonstrated that 
receiving devices can be operated rapidly enough by 
this far infrared radiation to control homing missiles. 
With the aid of a scanning device utilizing a rotating 
mirror, heat targets can be distinguished from alti¬ 
tudes up to 10,000 ft. 

The heat-homing missile, Felix, has been designed 
to operate in this range. Tests currently under way 
will show whether its reliability in the field will corre¬ 
spond to its laboratory indications. 
































GLOSSARY 


AAF. Army Air Forces. 

AFC. Automatic flight control. 

AGC. Automatic gain control. 

Altitude Signal. The radar signal returned to an airborne 
radar set by the ground or sea surface directly beneath the 
aircraft. 

Angle of Attack. The angle between a reference line fixed 
with respect to an airframe and the apparent flow line of the 
air through which it flies. 

Arrow Stability. The partial derivatives of yawing and 
pitching moments with respect to angles of attack in yaw 
and pitch. 

ASG. A specific airborne radar search set. 

Aspect Ratio. The ratio of span to mean chord of an airfoil; 
therefore, in wingless missiles such as Azon and Razon, the 
ratio of the bomb diameter to its mean length. 

Axis of Scan. In a scanning system, the axis about which in¬ 
formation as to the target location is collected and with 
reference to which target displacement is measured. 

Azon. A visually directed remotely controlled bomb radio con¬ 
trolled in AZimuth ONly. The AAF nomenclature for the 
1,000-lb version is V B-l; for the 2,000-lb version, it is V B-2. 

Back Coupling. A mechanical feedback link which displaces 
the scanning system through an angle which is a function of 
the rudder and/or elevator displacement. 

Ballodromic. Heading to hit; leading (from ballein, to hurl, to 
hit; dromos, course). 

Bat. A glide bomb of controlled flight steered by full-span 
trailing-edge wing flaps and a fixed empennage structure. 
Guidance is accomplished by a radar transmitter and re¬ 
ceiver on the missile. The bomb is steered to fly toward the 
direction from which the distinctive reflection is received. 

Beam-Receptor Character. The characteristic of an antenna 
array which gives it a maximum sensitivity in a single direc¬ 
tion, with continuously decreasing sensitivity with angular 
departures from this direction. 

Bore-Sighted. Aligned. The phrase is borrowed from gunnery 
usage where it means aligning the sight of a piece with its 
bore. In the present volume, it refers to the alignment of the 
axis of scan (supra) of the intelligence device of a missile 
with some appropriate reference axis such as the tangent to 
the flight path. 

Bore-Sight Error. Error in alignment of the axis of scan of 
the intelligence device of a missile. See Bore-Sighted. 

Brachydromic. Heading short; slanting to pass through the 
wake (from brachys, short; dromos, course). 

Canard Type of Head Controls. Aerodynamic control sur¬ 
faces placed at the nose of the fuselage of an airframe as con¬ 
trasted with the conventional aircraft empennage. 

Cardan-Mounted. Gimbal-mounted. 


Cassegrain Mirror. A plane mirror mounted between the 
surface of a spherical (or parabolic) mirror and its focus. The 
purpose is to project the image formed by the out of the in¬ 
cident rays. Named after Cassegrain the astronomer, who 
invented it. 

Chord. The dimension of an airfoil perpendicular to its span. 
The name derives from the chordal character of such a 
dimension with respect to the curved surfaces of the airfoil. 

Circle of Confusion. The smallest circle which can be re¬ 
solved by an optical system. 

Clamp Circuit. A bias-control circuit for the output stage of a 
video amplifier. 

Clinodromic. Heading at a constant lead angle (from klinein, 
to lean or to incline; dromos , course). 

Clinoscopic. Looking aslant, specifically sighting to lead a 
target (from klinein, to lean, to incline; skopos, target, aim). 

Compliance. The reciprocal of stiffness. 

Compressibility Burble Drag. The large and sudden in¬ 
crease in parasitic drag believed to be experienced by an air¬ 
foil as it passes through the velocity of sound. 

Control-Plane, adj. The qualifying term which describes 
the transmitting antenna on an aircraft which radiates the 
control signal by which a guided bomb is steered. 

Crab. An attachment to the Norden bombsight which super¬ 
poses on the field of view of its telescope an image of a falling 
bomb at its predicted point of impact. 

Crab Angle. The angle between the direction in which an air¬ 
craft is heading and its true course. 

Cross-Path Range. The direct range from a pursuing air¬ 
craft to its quarry when the former is flying a true pursuit 
curve. 

Dark Time. The time interval or portion of a time cycle when 
a photosensitive device (e.g., a phototube) is dark. 

Dead Band. The region near zero where an instrument sensi¬ 
tive to positive or negative values of a quantity gives no 
response. 

Dead Region. Dead band. 

Depression Angle. The angle measured downward from the 
horizontal to the axis of an airborne radar beam directed at 
a target. This is the complement of the incidence angle of 
the beam at the target plane. 

Dove Eye. The thermosensitive element used to control the 
heat-homing missile, Dove. 

Drag-Weight Ratio. The ratio of drag of an airframe to its 
total weight. 

Ears. Wind vanes mounted in the wind stream surrounding a 
missile and used to align a homing device (or television 
image) axis with the line of flight. 

Elevons. Wing flaps combining the functions of ELEVators 
and ailerONS. 


359 


360 


GLOSSARY 


EMR. Electro Mechanical Research, Inc. 

Felix. A heat-homing missile developed by the Massachusetts 
Institute of Technology. 

Flow Spoiler Type. A type of airframe control in which the 
smooth flow around an airfoil is interrupted or “spoiled” or 
so disturbed as to destroy in part the lift. 

Free Gyro. A gyroscope mounted in two (or more) gimbal 
rings so that its spin axis is free to maintain a fixed orienta¬ 
tion in space. 

Frequency Pulling. The tendency of a modulation stage to 
change the frequency of a direct-coupled master oscillator. 

Gating. The act of making a radar receiver operative for short 
intervals periodically spaced so that in effect a “gate” is 
opened for wanted signals and closed to others. 

Glide Bomb. A winged missile powered by gravity. The wing 
loading is so high that it is incapable of flight at the speeds 
of conventional bombardment aircraft. Such a missile must, 
therefore, be carried rather than be towed. 

Glide-Path Angle. The angle, measured from the horizontal, 
of inclination of the tangent to the glide path. 

Glide-Path Ratio. The ratio of lift to drag. This is the co¬ 
tangent of the glide-path angle. 

Glider (Bomb). A winged missile powered by gravity. The 
wing loading is sufficiently low so that it is capable of flight 
at the speeds of conventional bombardment aircraft. Such a 
missile may, therefore, be towed rather than be carried. 

GP Bomb. General purpose bomb. 

GSAP (Camera). Gun-sight aiming-point camera. An auto¬ 
matic motion picture camera put in operation by the pres¬ 
sure on machine-gun trigger. It was developed and used by 
the Services to record precision of aim in fixed-mount gun¬ 
nery. 

Head-End Tuner. In a radio receiver of the superheterodyne 
type, one or more tuned radio-frequency stages in advance 
of the first detector. 

Helicodromic. Heading along bent skew spiral. 

Hinge Moment. The moment tending to restore a control sur¬ 
face which has been displaced from a position of equilibrium 

Homing Intelligence. A signal received in a missile from its 
target which can be made to cause the missile to fly toward 
or to “home on” the target. 

Hunting. An oscillation about a mean or intended value as: a 
course, a desired frequency, a control-surface deflection, etc. 

Interlacing. A technique in television scanning wherein, if 
the lines are sequentially numbered, all the odd-numbered 
lines are scanned first following which all the even-numbered 
lines are scanned. 

Interrupter Controls. See “Flow Spoiler Type.” 

Intervalometer. An instrument associated with the bomb- 
sight which measures the time interval between the release 
of several bombs dropped in train. 

Isobars. Curves of constant acceleration loading of a pursuing 
aircraft flying a true pursuit curve. 


Jag. A device attached to the Norden bombsight which biases 
the synchronism adjustment of the sight to provide partial 
correction for systematic variation in time of fall of a Razon 
caused by application of control. 

Ladder Filter. A chain of T or 7r filters. 

Latitude. The range in brightness of a scene over which 
fidelity of response of a television pickup tube (or photo¬ 
graphic emulsion) is maintained. 

Lead Prediction. The act of directing a missile (or projectile) 
ahead of a moving target—leading in aim—to a predicted 
collision point. 

Leakance. A reciprocal of resistance. Conductance. 

Leveling Circuit. An r-c filter circuit used to level out fluc¬ 
tuations of a bias voltage. 

Lubber Mark. A mark on the casing of a compass which gives 
the heading of an aircraft or vessel carrying the compass. 

Mach Number. The ratio of the velocity of an aircraft to the 
velocity of sound at the location of the aircraft. 

Mimo. The Miniature IMage Orthicon, a television pickup 
tube. 

Mimo-Plane. The television receiving antenna located on the 
aircraft controlling Roc by means of television. 

Mimo-Roc. The television transmitting antenna on the Roc 
missile. 

MIT. Massachusetts Institute of Technology. 

Mosaic Cathode. The photo cathode in the iconoscope char¬ 
acterized by being made up of discrete particles like a 
mosaic. 

Multipaths. The several paths by which, due to reflections, a 
radiated signal, particularly a television signal, may reach 
the receiving antenna from the transmitter. 

NACA. National Advisory Committee for Aeronautics. 

NBS. National Bureau of Standards. 

Nodding Motion. A method of inclining a homing receiver so 
that it can “look chordally at” the target across the tra¬ 
jectory during the early phases of the flight of a homing 
missile when the attitude of the missile is relatively hor¬ 
izontal. 

Orchard Heater. An oil-fired heater used to ward off frost in 
citrus orchards. 

Parabolic Thinking. Nodding motion. 

Pelican. A radar-homing glide bomb equipped with a radar 
receiver and which automatically homes on radar reflections 
from an airborne radar set mounted in the aircraft which 
releases the missile. It is also characterized by full-span, 
trailing-edge wing flaps and a fixed empennage structure so 
that it flies with substantially constant angle of attack. 

Phantasmagoria. A simulated guided-missile system by which 
one or more of the motions of the missile may be dynami¬ 
cally studied at a reduced scale in space and an expanded 
scale in time. 

Phoronomy. Science of relative motion of moving bodies with 
respect to each other. 



GLOSSARY 


361 


Phugoid. A path of flight; specifically, the path of a glider 
having longitudinal stability and having its control surfaces 
fixed. 

Pick-Off. A means of having a sensitive instrument, as for 
example a gyroscope, actuate a control system or an element 
thereof. 

Pitch, v. To rotate, usually through a small angle, about a 
horizontal axis called the pitch axis, which is perpendicular 
to the longitudinal axis of an aircraft or ship. n. The act 
of pitching, adj. Of, or pertaining to, a pitching motion. 

Pitch Gyro. A gyroscope so mounted in an aircraft as to be 
sensitive to pitch. 

Pitch-Input Link. A link in a servomechanism or simulator 
whereby the motion of pitch is inserted into the system for 
control or study. 

Presentation Tube. A cathode-ray tube for the presentation 
of a signal such as a radar signal or a television picture. 

Proportional Control. Control in which the action to cor¬ 
rect an error is made proportional to that error. 

Proximity Parameter. A parameter which expresses the 
distance between an aircraft, flying a pursuit course, and its 
quarry when the pursuer passes abeam of the quarry. 

Pull-Out. The act of changing from a power dive to level (or 
climbing) flight. 

Radar Dish. The dish-shaped reflector behind a radar 
antenna. 

Radar Width. In a radar homing set, the range of error within 
which the response is proportional to the error in heading. 

Range-Tracking Element. An element in a radar set which 
measures range and its time derivative. By means of the 
latter, a range gate is actuated at the predicted instant of 
signal reception. 

Rapid Incidence Adjustment. A short-period highly damped 
oscillation of an aircraft about its pitch axis following a 
sudden change in altitude. 

Raster. A system of luminescent lines traced on the phosphor 
of a cathode-ray tube by motion of the cathode-ray beam. 
The changes of brightness in the lines produce a picture as a 
television picture or a radar map. This word is of German 
origin and is used in particular in television. In the present 
volume, its use is limited to a system of parallel lines for the 
production of a television picture. 

Rate Gyro. A gyroscope with a single gimbal mounting such 
that angular rates of rotation about an axis perpendicular to 
the axis of the gimbal mounting produce measurable pre- 
cessional forces. 

Razon. A high-angle radio-controlled bomb visually guided in 
Range and AZimuth ONly. 

RCA. Radio Corporation of America. 

Redundance. The property of an equation which permits a 
plurality of solutions; therefore, in a mechanical system 
describable by such an equation, the property which permits 
a plurality of modes of action. 


RHB. Radar homing bomb; specifically, one in which the radar 
transmitter is not located in the missile, e.g., Pelican. 

Ringing. Sustained oscillation which occurs in a high-gain 
amplifier inadequately stabilized. 

Roc. A medium angle controlled missile characterized by 
(1) four symmetrical wings with full-span wing flaps and 
four fixed fins or, (2) a single cylindrical wing capable of 
being rocked about the yaw and pitch axes and a fixed 
cylindrical empennage. 

Roll. v. To rotate, usually through a small angle, about an 
axis substantially collinear with the longitudinal axis of an 
aircraft, or ship. n. The act of rolling, adj. Of, or per¬ 
taining to, rolling motion. 

Run-Out Time. The time of travel, from neutral to full ex¬ 
tension, of the control surfaces of an airframe. 

Sailplane. A glider capable of soaring. 

Sap Bomb. Semi-armor-piercing bomb. 

Saturation Angle. Radar width. 

Scopodromic. Heading as looking; homing (from skopos, tar¬ 
get, aim; dromos, course). 

Servo-Link. A mechanical power amplifier by which signals 
at a low power level are made to operate control surfaces re¬ 
quiring relatively large power inputs; e.g., a relay and motor- 
driven actuator. 

Servomechanism. The feedback loop of a servo system exclu¬ 
sive of the missile itself. 

Servo System. A closed feedback loop comprising the intel¬ 
ligence device, automatic pilot if any, the amplifying link, 
and the missile itself. 

Shading. The appearance of dark areas in a received television 
picture which sometimes cover the entire screen. 

Sink. A point or element in a system where energy is dissipated 
or otherwise removed from the system. 

Slant Range. The range to the target along a line of sight. 

Spoiler. A surface which is projected into the wind stream 
surrounding an airfoil and “spoils” or interrupts the air flow 
reducing the lift. 

Squib. An electrically ignited charge of explosive contained in 
a cylinder used to actuate a process remotely, e.g., to release 
a parachute. 

SRB. Send-Receive Bomb, a radar homing bomb in which the 
transmitter as well as the receiver is carried in the missile, 
e.g., Bat. 

Synchro. A dynamo electric machine having a polyphone 
stator and a single-phase rotor. 

Tearing. The destruction of a received television picture due 
to interference or malefaction characterized by the appear¬ 
ance of tearing. 

Time Off. The time interval during which an intermittently 
energized control element is de-energized. 

Time On. The time interval during which an intermittently 
energized control element is energized. 






362 


GLOSSARY 


Tomodromic. Heading to cut; intersecting (from tomos , a cut¬ 
ting; dromoSy course). 

Track in Range. To adjust the gate in a radar set so that it 
opens at the correct instant to accept the signal from a tar¬ 
get of changing range from the radar. 

Trail Angle. The angle between a line joining the point of 
impact of a bomb and the releasing aircraft at the instant of 
impact, and the vertical, the aircraft having performed level 
unaccelerated flight from the point of release. 

Transfer Function. The function relating the output of a 
closed-cycle servo system to its error. 

Trembling. Shaking, vibrating, oscillating. 

Tumble. The act performed by a 2-frame free gyroscope when 
both frames become co-planar. Under these circumstances, 
the gyro wheel rotates about a diameter as well as about its 
polar axis. Control is lost. 

Umbilical Switch. A switch located in the cable connecting a 
missile with the aircraft during carriage. This switch is 
actuated upon release of the missile and the pulling out of 
the cable. 


“Venetian Blind” Structure. A form of electron multiplier 
structure in the image orthicon pickup tube; so-called on 
account of its cross-sectional resemblance to a Venetian 
blind. 

Water-Vapor Window. The region in the electromagnetic 
spectrum between 8.5- and 13 -m wavelength radiations 
where water vapor is transparent. 

Weathercock Stability. The partial derivatives of yawing 
and pitching moments with respect to angles of attack in 
yaw and pitch. 

Wing-Incidence Angle. The angle between the airstream 
incident at the leading edge of a wing and the direction 
which such an airstream would have were it to produce zero 
lifts. 

Yaw. v. To rotate, usually through a small angle, about an axis 
perpendicular to the pitch and roll axes of an aircraft or 
ship. For an aircraft in normal level flight, the yaw axis is 
vertical; in a missile, however, it adopts an inclination as the 
missile noses over on its trajectory. With a high-angle bomb 
dropped from great height, for example, the yaw axis is 
nearly horizontal at impact, n. the act of yawing, adj. Of, 
or pertaining to, yawing motion. 




BIBLIOGRAPHY 


Numbers such as Div. 5-210-MI indicate that the document listed has been microfilmed and that its title appears in 
the microfilm index printed in a separate volume. For access to the index volume and to the microfilm, consult the Army 
or Navy agency listed on the reverse of the half-title page. 


Chapter 1 

1. Aerodonetics, F. W. Lanchester, Constable, London, 
1906, Chapter 9. 

2. Pelican Ballistics , AMP Memo 111-1M, SRG-C 561, 
SRG-P 155. 

3. Wind-Tunnel Tests of a Target Glider, Robert A. Conway, 
Memorandum Report, NACA, Oct. 15, 1943. 

4. Analysis of the Longitudinal Stability of Homing Glide 

Bombs with Application to Navy SWOD Mark 7 and Mark 
9, Harold K. Skramstad, Service Projects AC-1, NO-115, 
and others, NBS, Oct. 15, 1945. Div. 5-210-MI 

5. Nomenclature for Aeronautics , Report 474, NACA, 1941. 

6. An Automatic Pilot for Homing Glide Bombs , John A. 
Hart, Service Project AC-1, NO-115, and others, NBS. 

Div. 5-210-M3 

7. Radio Set RHB , Elisabeth M. Lyman (editor), OEMsr- 

262, Reports 508-1, 508-2 and 508-3, Radiation Labora¬ 
tory, MIT, Jan. 17, 1944. Div. 5-320-MI 

8. Developments in Radar Homing Missiles for the Period 

March 1, 1944 to May 1, 1945, R. A. Lamm, A. J. 
McLennan, and P. R. Stout, OSRD 5549, OEMsr-240, 
Service Projects AC-36, NO-115, and others, MIT, June 
1, 1945. Div. 5-320-M2 

9. Final Report on Developments in Radar Homing Missiles, 

R. A. Lamm, P. R. Stout, and H. R. Toporeck, OSRD 
6413, OEMsr-240, Service Projects AC-36, NO-115, and 
others, MIT, Nov. 30, 1945. Div. 5-320-M3 

10. The Development of Servo-Control Mechanisms for Homing 

Aero-Missiles, Emmett C. Bailey and Wesley G. Span- 
genberg, Service Projects AC-1, NO-115, and others, 
t^BS. Div. 5-350-M4 

11. Representation of Longitudinal Stability and Control of 

Homing Glide Bombs by an Electro-Mechanical Model, 
Harold K. Skramstad, Service Projects AC-1, NO-115, 
and others, NBS, Oct. 26, 1945. Div. 5-210-M2 

Chapter 2 

1. Aerodynamics of Aircraft Bombs, Hugh L. Dryden, NBS, 

Feb. 28, 1927. Div. 5-100-MI 

2. Progress Report No. 11, Early Razon Experiments, OSRD 
4557, Gulf Research and Development Company, Sept. 1, 
1944. 

3. Secret letter from W. J. Richards to Director General 
British Air Commission, Attention: Air Vice Marshal 
Mansell, dated June 10,1944, OSRD Liaison Office Docu¬ 
ment WA-2356-11A. 

4. Seventh Progress Report on the Development of a Dirigible 

Bomb, J. F. Hutzenlaub, OSRD 1704, OEMsr-240, MIT, 
Apr. 15, 1945. Div. 5-231-M2 


5. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, OSRD 1397, NDCrc-183, Gulf 
Research and Development Company, Mar. 15, 1943. 

Div. 5-232.1-MI 

6. Experimental Investigations in Connection with High- 

Angle Dirigible Bombs. Early Razon Experiments, OSRD 
4557, NDCrc-183, Gulf Research and Development Com¬ 
pany, Sept. 1, 1944. Div. 5-232.2-MI 

7. Progress Report, OSRD 6187, Gulf Research and De¬ 
velopment Company, Apr. 1, 1945. 

8. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, VB-3 Ballistic Data. OSRD 6221, 
NDCrc-183, Service Project AC-36, Gulf Research and 
Development Company, Sept. 1, 1945. Div. 5-232.2-M2 

9. Colored Flares, G. Albert Hill and R. G. Clarke, OSRD 
4408, OEMsr-321 Wesleyan University, Nov. 25, 1944. 

Div. 11-202.12-MI 

10. Colored Flares, G. Albert Hill and R. G. Clarke, OSRD 
4408-A, OEMsr-321, Wesleyan University, Feb. 10, 
1945. Div. 5-333-MI 

Moving Picture Camera Installation in Bombs to Study 
Flight Characteristics of Bombs with Special Reference to 
Their Yaw and Spin Motions, OSRD 106, NDCrc-183, 
Gulf Research and Development Company, Sept. 1,1941. 

Div. 5-421.1-MI 

Experimental Investigations in Connection with the Flight 
Characteristics of Bombs, T. B. Pepper, OSRD 232, 
NDCrc-183, Gulf Research and Development Company, 
Apr. 1, 1942. Div. 5-231-MI 

First Report, Air Forces Evaluation Board, Pacific Ocean 
Area, OSRD Liaison Office Document LOGA J-5809, 
Aug. 28, 1944. 

The High-Angle Dirigible Bomb Project, OSRD 6379, 
NDCrc-183, Service Project AC-36, Gulf Research and 
Development Company, Oct. 31, 1945. Div. 5-231-M5 

Chapter 3 

1. First Progress Report: Far Infrared Detecting Devices, 
OSRD 1157, Harvard University, Dec. 31, 1942. 

Div. 16-310.2-MI 

2. “Measurement of Thermal Radiation at Microwave Fre¬ 
quencies,” R. H. Dickie, Review of Scientific Instruments, 
Vol. 17, No. 7, July 1946. 

3. Felix, A High-Angle, Heat-Homing Bomb, Alan C. Bemis, 

OSRD 6416, NDCrc-180, Service Project AC-36, MIT, 
Oct. 31, 1945. Div. 5-240-M3 

Felix, Alan C. Bemis, OSRD 4039, NDCrc-180, Service 
Project AC-36, MIT, June 1, 1944. Div. 5-240-MI 


11 . 

12 . 

13. 

14. 



363 


364 


BIBLIOGRAPHY 


5. “Theory of Servomechanisms,” H. L. Hazen, Journal of 
Franklin Institute, September 1934. 

6. On Black-White Control with Position Correlation, K. 
Klatter and I. Lotz, U & M 1326. 

7. On the Influence of Input Lag, Scholz, U & M 1327. 

8. On Black-White Control with Velocity Correlation, K. 
Klatter and Hodnapp, U & M 1328. 

9. Application to Pitch Control, I. Lotz and Meissinger, 
U & M 1329. 

10. Summary, I. Lotz, U & M 1330. 

11. Regelkreisschwingungen hei der A uf-Zu-Regelung ( Schwarz- 
Weiss-Steuerung) . H. Greinel and H. Leisegang, DFS Re¬ 
port. 

12. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, The Photo-Electric Target-Seeking 
Bomb, OSRD 3422, NDCrc-183, Gulf Research and 
Development Company, Jan. 1, 1944. Div. 5-231-M3 

13. Film Type Bolometers. Charles E. Aiken, William H. Car¬ 

ter, Jr., and F. S. Phillips, OSRD 5530, OEMsr-921, 
Service Project AC-36, Electro-Mechanical Research, 
Inc. Aug. 13, 1945. Div. 5-332-M3 

14. Felix, A High-Angle, Heat-Homing Bomb, Alan C. Bemis, 

OSRD 6416, NDCrc-180, Service Project AC-36, MIT, 
Oct. 31, 1945, Appendix A. Div. 5-240-M3 

15. Metal-Strip Bolometers, Bruce Billings, Edgar Barr, and 

Lewis Hyde, OSRD 6397, OEMsr-1317, NOrd-6013, 
Service Project NO-257 and AC-36, Polaroid Corpora¬ 
tion, March 1946. Div. 5-332-M4 

16. The Heat-Homing Eye. OEMsr-1317, Service Projects 
AC-36, NO-257, Polaroid Corporation, Aug. 15, 1945. 

Div. 5-330-MI 

17. Report on Project Dove in Fourteen Volumes, NOrd- 
6013 Polaroid Corp., March 1946. 

18. A Survey of Targets for Felix, Alan C. Bemis OSRD 6661, 
NDCrc-180, Service Project AC-36, MIT, Oct. 31, 1945. 

Div. 5-240-M2 

Chapter 4 

1. Phorodromics, W. B. Klemperer, Report SM-3536, 
Douglas Aircraft Company, Inc., July 17, 1942. 

Div. 5-100-M2 

2. Developments in Radar Homing Missiles for the Period 

March 1, 1944 to May 1, 1945, R. A. Lamm, A. J. Mc¬ 
Lennan, and P. R. Stout, OSRD 5549, OEMsr-240, 
Service Projects AC-36, NO-115, and others, MIT, 
June 1, 1945. Div. 5-320-M2 

3. “Flight Characteristics,” W. B. Klemperer, Echodromics, 

Report SM-3745, Douglas Aircraft Company, Inc., Dec. 
31, 1942, Chapter 1. Div. 5-100-M3 

4. Operational Circuit Analysis, Vannevar Bush, John Wiley 
and Sons, 1924. 

5. Report DF-71370, General Electric Co., February 1946. 
(Sequel XIV contains a summary in which an equation in 
0, 0, <f>, <f>, <j> is quoted.) 


6. “Design of Flight Test Bird,” W. B. Klemperer, Echo¬ 
dromics, OSRD 3654, Report SM-3745, Douglas Aircraft 
Company, Inc., Dec. 1,1943, Chapter 4. Div. 5-221-MI 

7. W ind Tunnel Tests, Model Roc-1,” W. B. Klemperer 

and R. S. Shevell, Echodromics, Report SM-3745, 
Douglas Aircraft Company, Inc., Oct. 15, 1943, Chapter 
5 - Div. 5-222-MI 

8. Control Amplifiers for Guided Missiles, W. S. Leitch, 

OSRD 6417, OEMsr-1002, Service Project AC-36, Ben- 
dix Aviation Corp., Jan. 30, 1946. Div. 5-223-M2 

9. Functional Tests, Brood 1-A, Test Birds with Quadrant 

Eyes,” W. B. Klemperer, Project Roc, OSRD 3937, Re¬ 
port SM-3745, Douglas Aircraft Company, Inc., Apr. 18, 
1944, Chapter 6. Div. 5-222-M2 

10. “Brood 1-B, Test Birds with Fairchild Eyes,” W. B. 

Klemperer, Project Roc, OSRD 4213, Report SM-3745, 
Douglas Aircraft Company, Inc., July 25, 1944, Chapter 
7 - Div. 5-222-M3 

11. “Functional Test, Brood 1-B, Test Birds with Fairchild 
Eyes,” Echodromics, OSRD 4213, Douglas Aircraft Com¬ 
pany, Inc., July 25, 1944, Chapter 7. 

12. Design and Development of the ROC 00-1000 Serial No. 1 

to No. 21, OSRD 6330, OEMsr-327, Service Project 
AC-36, Report SM-8349, Douglas Aircraft Company, 
Inc., June 15, 1945. Div. 5-221-M2 

13. Design Modifications of the ROC 00-1000 Serial No. 21 

through No. 55, OSRD 6609 OEMsr-327, Service Project 
AC-36, Report RL-8384, Douglas Aircraft Company, 
Inc., Jan. 25, 1946. Div. 5-221-M3 

14. Tests of a %-Scale Model of the NDRC ROC 00-1000 Mis¬ 
sile, Kenneth M. Hughes, Memorandum A-4K10, Ames 
Aeronautical Laboratory, NAGA, Nov. 10, 1944. 

Div. 5-222-M5 

15. Experimental Study of Control Techniques for Television 

ROC, with Appendices by Dr. P. Mertz, Division 5, Dr. 
Ruth M. Peters Section 7.4, and Mr. Lobe Julie, Columbia 
University, E. H. Pier, OSRD 6562, OEMsr-327, Service 
Project AC-36, Douglas Aircraft Company, Inc., Feb. 15, 
1946. Div. 5-223-M3 

16. Flight Tests with ROC 00-1000V, Drops V-l through V-21, 
OSRD 6414, OEMsr-327, Service Project AC-36, Douglas 
Aircraft Company, Inc., Jan. 15, 1946. Div. 5-222-M6 

17. Flight Tests of the ROC 00-1000T, Drops T-l through T-10, 
OSRD 6610, OEMsr-327, Service Project AC-36, Douglas 
Aircraft Company, Inc., Feb. 25, 1946. Div. 5-222-M7 

Chapter 5 

1. “Television Equipment for Guided Missiles,” Charles J. 
Marshall and Leonhard Katz, Proceedings of the Institute 
of Radio Engineers, Vol. 34, No. 6, June 1946. 

2. “Television Studio Design,” R. M. Morris and R. E. 
Shelby, RCA Review, July 1937. 

3. “A Study of the Television Image Characteristics,” E. 
W. Engstrom, Proceedings of the Institute of Radio Engi¬ 
neers, December 1933. 





BIBLIOGRAPHY 


365 


4. “The Iconoscope—A Modern Version of the Electric Eye,” 
V. K. Zworykin, Proceedings of the Institute of Radio 
Engineers, January 1934. 

5. The Image Orthicon, Albert Rose, Paul K. Weimer, and 

Harold B. Law, OSRD 3724, OEMsr-441, RCA Labora¬ 
tories. February, 1944. Div. 5-421.21-MI 

6. Conversion Unit Comparison, Paul S. Hendricks, OSRD 
5105, OEMsr-1191, Service Projects AC-36 and NA-116, 
Columbia Broadcasting System, Inc., Mar. 15, 1945. 

Div. 5-421-MI 

7. Military Television Equipment, Walter L. Lawrence, 

OSRD 6415, OEMsr-441, Service Projects AC-36, NA- 
190, and NS-132, RCA Victor Division of Radio Corpora¬ 
tion of America, Jan. 29, 1946. Div. 5-400-M3 

8. Block Simplification and Improvement W. J. Poch and A. 

H. Turner, OSRD 6394, OEMsr-441, Service Projects 
AC-36, NA-190, and NS-132, RCA Victor Division of 
Radio Corporation of America, Nov. 20, 1945. 

Div. 5-420-M5 

9. General Block Improvements, T. L. Gottier, OSRD 4222, 
OEMsr-441, RCA Laboratories, Apr. 21, 1944. 

Div. 5-420-M2 

10. Development of Block Equipment, OSRD 2032, RCA 
Victor Division of Radio Corporation of America. Aug. 1, 
1943. 

11. Improved Sensitivity of Television Equipment, Ray D. 
Kell, RCA Laboratories, OSRD 3174, OEMsr-441, Aug. 

I, 1943. Div. 5-421.1-M3 

12. Television Camera Employing Schmidt Optics, E. Dudley 
Goodale, L. R. Moffett, and W. L. States, OEMsr-515, 
National Broadcasting Company, Inc., Sept. 20, 1942. 

Div. 5-421.1-M2 

13. Sealed Beam Dissector, Madison Cawein, OSRD 5215, 
OEMsr-1093 and OEMsr-690, Farnsworth Television 
and Radio Corporation, Apr. 5, 1945. Div. 5-421.22-M2 

14. Vericon Television Equipment, OSRD 5289, OEMsr-187> 
Remington Rand Inc., Electronic Division, Feb. 1, 1945. 

Div. 5-421.23-MI 

15. Development of a Television Pickup Tube of Greater 

Sensitivity, Albert Rose, OSRD 6393, OEMsr-441, 
Service Projects AC-36, NA-190, and NS-132, RCA 
Victor Division of Radio Corporation of America, Oct. 
31, 1945. Div. 5-421.21-M2 

16. Frequency Modulation for the Transmission of Television 

Signals, Robert W. Clark, Robin D. Compton, and 
Vernon J. Duke, OEMsr-513, National Broadcasting 
Company, Inc., Sept. 20, 1942. Div. 5-410-MI 

17. AM-FM Comparison, Robert W. Clark, OSRD 1988, 

OEMsr-441, National Broadcasting Company, Inc., 
June 30, 1943. Div. 5-410-M2 

18. Final Report on Block X Equipment, R. C. Moore, OSRD 

4579, OEMsr-1159, Philco Radio and Television Cor¬ 
poration, Oct. 17, 1944. Div. 5-420-M3 

19. Development of Block Equipment, H. N. Kozanowski, 
OSRD 2031, OEMsr-441, RCA Victor Division of Radio 
Corporation of America, Aug. 15, 1943. Div. 5-420-MI 


20. Block XVIII Television Equipment, H. B. Fancher, OSRD 
4670, OEMsr-1172, General Electric Co., June 1, 1944. 

Div. 5-420-M4 

21. Final Report on Development of High-Sensitivity Airborne 

Television Transmitting Equipment, A. V. Bedford, OSRD 
6655, OEMsr-441, Service Projects AC-36, NA-190, and 
NS-132, RCA Victor Division of Radio Corporation of 
America, Feb. 6, 1946. Div. 5-400-M4 

22. Field Service Report, Alfred E. Jackson, OSRD 1989, 

OEMsr-441, National Broadcasting Company, Inc., 
June 30, 1943. Div. 5-400-MI 

23. Block Equipment Used in Glider Bombs, Eglin Field, 

Florida, Robert W. Clark, Alfred R. Jackson, and R. R. 
Davis, OSRD 1990, National Broadcasting Company, 
Inc. Div. 5-311-M4 

24. Field Service, Ray D. Kell, OSRD 4481, OEMsr-441, 

RCA Laboratories, Aug. 1, 1944. Div. 5-400-M2 

25. Television-Equipped Radio-Controlled Bomb VI A. V. 
Loughren and John A. Hansen, Report D3-288, Hazeltine 
Service Corporation, Sept. 18, 1942. Div. 5-311-MI 

26. Television-Equipped Radio-Controlled Bomb IX, A. V. 
Loughren and John A. Rado, Report D3-291, Hazeltine 
Service Corporation, Sept. 22, 1942. Div. 5-311-M2 

27. Final Report on NDRC Dissector-Camera Project, Madison 
Cawein, OSRD 5201, OEMsr-620 and OEMsr-1093, 
Service Projects AC-36 and NA-116, Farnsworth Tele¬ 
vision and Radio Corporation, Feb. 1, 1945. 

Div. 5-421.22-MI 

28. Television-Equipped High-Angle Projectiles, H. P. See, 

OEMsr-615, National Broadcasting Company, Inc., Jan. 
23, 1943. Div. 5-311-M3 

29. Tests of Television-Equipped High-Angle Bombs, Alfred E. 

Jackson, OSRD 4296, OEMsr-441, National Broadcast¬ 
ing Company, Inc., May 1944. Div. 5-231-M4 

Chapter 6 

1. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, OSRD 1397, NDCrc-183, Gulf 
Research and Development Company, Mar. 15, 1943. 

Div. 5-232.1-MI 

2. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, OSRD 1785, NDCrc-183, Gulf 
Research and Development Company, May 15, 1943. 

Div. 5-232.1-M2 

3. The Azon Bomb, OSRD 3086, NDCrc-183, Gulf Research 
and Development Company, Oct. 15, 1943. 

Div. 5-232.1-M3 

4. VB-1 Tail Assembly, OSRD 5288, OEMsr-1081, Service 

Project AC-36, Union Switch and Signal Construction 
Company, June 15, 1945. Div. 5-232.1-M4 

5. Radio Link Receiver and Crystal-Controlled Superhetero¬ 
dyne Receivers for Azon and Razon, A. W. Friend, OSRD 
4139, OEMsr-240 and OEMsr-1195, NDCrc-180, MIT 
and Harvey Radio Laboratories, July 15, 1944. 

Div. 5-232.21-MI 





366 


BIBLIOGRAPHY 


6. Radio Link Receiver Equipment, Joseph C. Tellicr, OSRD 
4327, OEMsr-1314, Research Division, Phileo Radio and 
Television Corporation, September 1944. 

Div. 5-232.21-M2 

7. Design of Single Radio Band Interference Reducing Sys¬ 
tems for Radio Control Purposes , OSRD 3229, OEMsr- 
094, Hammond Research Corporation, Nov. 30, 1943. 

Div. 5-310-MI 

8. Improved Radio Control Equipment for Target Airplanes, 

Richard C. Webb, Report 1)3-340, OEMsr-278, Purdue 
University, Dec. 1, 1942. Div. 5-312-MI 

Chapter 7 

1. Fundamental Theory of Servomechanisms, L. A. MacColl, 
D. Van Nostrand Company, Inc., 1945. 

2. Development of a Servo Control System for Guided Missiles, 

OSRD 5514, OEMsr-1013, Service Projects AC-36 and 
NO-115, Servomechanisms Laboratory, MIT, July 2, 
1945. Div. 5-350-M2 

3. Development of a Servo-Control System for Guided Missiles, 

OSRD 6418, OEMsr-1013, Service Projects AC-36 and 
NO-115, Servomechanisms Laboratory, MIT, Oct. 31, 
1945. Div. 5-350-M3 

4. The Analysis and Synthesis of Servomechanisms, A. C. 
Hall, The Technology Press, MIT, May 1943. See also 

references cited. Div. 5-350-MI 

• 

5. “Regeneration Theory," H. Nyquist, Bell System Tech¬ 

nical Journal, January 1932. 

Chapter 8 

1. The Life of William Thomson, Sylvanus P. Thompson, 
MacMillan and Company, Ltd., London, 1910. 

2. Survey of Facilities in Germany for the Development of 
Guided Missiles, Part V, Report of Alson Mission 
WBK-293, Sept. 15, 1945. 

3. Instrumentation of Homing Glide Missiles, Patterson, 
Lander, Collett, and Hart, Memorandum to Division 5 
from NBS. 

4. Data Recording Camera, Joseph L. Boone, OSRD 4255, 
OEMsr-978, Eastman Kodak Company, Sept. 9, 1944. 

Div. 5-222-M4 

5. Motion Picture Camera Installation in Bombs to Study 

Flight Characteristics with Special Reference to Their Yaw 
and Spin Motions, T. B. Pepper, OSRD 106, NDCrc-183, 
Gulf Research and Development Company, Sept. 1, 
1941. Div. 5-421.1-MI 

6. War Department Technical Manual TM 11-434D. 

7. “Functional Tests, Brood 1-A, Phototheodolite Tech¬ 
niques," W. B. Klemperer, Project Roc, OSRD 3937, Re¬ 
port SM-3745, Douglas Aircraft Company, Inc., Apr. 18, 
1944, Chapter 6, Appendix XXIV. Div. 5-222-M2 

8. Bomb Trajectory Camera , NDCrc-183, Service Project 

AC-36, Gulf Research and Development Company, Oct. 
31, 1945. Div. 5-234-Ml 


Chapter 9 

1. Experimental Investigations in Connection with High- 
Angle Dirigible Bomb, The Photo-Electric Target Seeking 
Bomb, OSRD 3422, NDCrc-183, Gulf Research and De¬ 
velopment Company, Jan. 1, 1944. Div. 5-231-M3 

2. “Design of Flight Test Birds," W. B. Klemperer, Echo- 

dromics, OSRD 3654, SM-3745, Douglas Aircraft Com¬ 
pany, Inc., Dec. 1, 1943. Div. 5-221-MI 

3. The Fairchild Wide-Angle Scanner, OSRD 4675, OEMsr- 

1182, Fairchild Camera and Instrument Corporation, 
May 24, 1944. Div. 5-331-M2 

4. An Anti-Rotational Stabilizer, H. G. Doll, G. K. Miller, 

and Charles B. Aiken, OSRD 6202, OEMsr-921, Service 
Project AC-36, Electro-Mechanical Research, Inc., Sept. 
15, 1945. Div. 5-360-M2 

5. A To-and-Fro Scanning Device Giving Right-Left Control, 

Charles B. Aiken, OSRD 3323, OEMsr-921, Service Pro¬ 
ject AC-36, Electro-Mechanical Research, Inc., Jan. 28, 
1944. Div. 5-331-MI 

6. Analysis of Bolometer Response to Modulated Energy, 
Charles B. Aiken, OSRD 3205, OEMsr-291, Service Pro¬ 
ject AC-36, Electro-Mechanical Research, Inc., Dec. 22, 

1943. Div. 5-332-M1 

7. The Frequency Characteristics of Bolometers, Charles B. 
Aiken, OSRD 4580, OEMsr-921, Service Project AC-36, 
Electro-Mechanical Research, Inc., Nov. 17, 1944. 

Div. 5-332-M2 

8. Heat-Sensitive Devices Employing To-and-Fro Scanning, 
Charles B. Aiken, M. LeBourg, and W. C. Welz, OSRD 
6175, OEMsr-921, Service Project AC-36, Electro- 
Mechanical Research, Inc., Sept. 20, 1945. 

Div. 5-331-M3 

9. Project Beetle, W. S. Leitch, OEMsr-1002, Bendix 

Aviation Corporation, Oct. 13, 1943. Div. 5-223-MI 

10. Present Status of the Birds Eye Bomb, B. F. Skinner, K. B. 

Breland, and Norman Guttman, General Mills, Inc. 
Feb. 1, 1943. Div. 5-340-MI 

11. Characteristics of the Response of the Vacuum System Unit, 

Philip Christopherson, OEMsr-1068, General Mills, Inc. 
Jan. 7, 1944. Div. 5-340-M2 

12. Final Report on Organic Homing Systems, B. F. Skinner, 
E. E. Kuiphal, and others, General Mills, Inc., Feb. 21, 

1944. Div. 5-340-M3 

Chapter 10 

1. “A Continuous Integraph,” Vannevar Bush, F. D. Gage, 
and H. R. Stewart, Journal of Franklin Institute, 203, 
63-84, (1927). 

2. “A New Type of Differential Analyzer," Vannevar Bush 
and S. H. Caldwell. Journal of Franklin Institute, 240, 
255-326, (1945). 

3. “Oscillographic Solution of Electro-Mechanical Systems," 
C. A. Nickle, Transactions of the American Institute of 
Electrical Engineers, 44, 844-855, (1925). 




BIBLIOGRAPHY 


367 


4. Experimental Investigations in Connection with High- 
Angle Dirigible Bombs, The Photoelectric Target-Seeking 
Bomb, OSRD 3422, NDCrc-183, Gulf Research and 
Development Company. Jan. 1, 1944. Div. 5-231-M3 

5. Felix, Alan C. Bemis, OSRD 4039, NDCrc-180, Service 
Project AC-36, MIT, June 1, 1944. Div. 5-240-MI 

6. Representation of Longitudinal Statrility and Control of 

Homing Glide Bombs by an Electro-Mechanical Model, 
Harold K. Skramstad, Service Projects AC-1, NO-115, 
and others, NBS, Oct. 26, 1945. Div. 5-210-M2 

7. Development of a Servo Control System for Guided Missiles, 

OSRD 5514, OEMsr-1013, Service Projects AC-36 and 
NO-115, MIT, July 2, 1945. Div. 5-350-M2 

8. Final Report to Section 7.2, OEMsr-1237, Columbia Uni¬ 
versity. 

9. The Azon-Razon Bombing Trainer: Preliminary Technical 
Manual, Division 5, NDRC, Aug. 15, 1945. 

Div. 5-233-M1 

10. Theory and Operation of Television Mimo Bomb Simulator, 

A. Costa de Beauregard, George Bartholomei, and others, 
OSRD 6396, OEMsr-1493, Service Project AC-36, Spe¬ 
cialties, Inc., Dec. 14, 1945. Div. 5-224-MI 

11. Experimental Study of Control Techniques for Television 
Roc, E. H. Pier, OSRD 6562, OEMsr-327, Service Project 
AC-36, Douglas Aircraft Company, Feb. 25, 1946. 

Div. 5-223-M3 

Chapter 11 

1. Final Report on Engineering Activities far Division 5, 
NDRC, Wallace H. Nichols, OEMsr-240, Service Project 
AC-36, MIT, Oct. 1, 1945. Div. 5-10-MI 

Chapter 12 

1. Aerodynamic Theory, William Frederick Durand, Julius 
Springer, Berlin, Vol. V, 1936. 

la. Ibid., Division N. 

2. The Analysis and Synthesis of Iyinear Servomechanisms, 

A. C. Hall, MIT, May 1943. Div. 5-350-MI 

3. The Theoretical Lateral Motions of an Automatically Con¬ 
trolled Airplane Subjected to a Yawing Moment Disturb¬ 
ance, F. H. Imlay, Restricted Technical Note 809, 
NACA. 

Appendix B 

1. Analysis of the Longitudinal Stalrility of Homing Glide 
Bombs with Applications to Navy SWOD Mark 7 and 
Mark 9, Harold K. Skramstad, Service Projects AC-1, 
NO-115, and others, NBS, Oct. 15,1945. Div. 5-210-M1 

2. Report on Wind Tunnel Tests on a 0.6 Scale Model of the 
Bureau of Standards Design NDRC Glide-Bomb, GALCIT 
Report 3.50, Mar. 3, 1942. 


3. The Effect of Lateral Controls in Producing Motion of an 
Airplane as Computed from Wind-Tunnel Data, Fred E. 
Weick and Robert T. Jones, Report 570, NACA, 1936. 

4. An Analysis of Lateral Stability in Power-Off Flight with 
Charts for Use in Design, Charles H. Zimmerman, Report 
589, NACA, 1937. 

5. An Automatic Pilot for Homing Glide Bombs, John A. 
Hart, Service Projects AC-1, NO-115, and others NBS. 

Div. 5-210-M3 

6. The Development of Servo-Control Mechanisms far Homing 
Aero-Missiles, Emmett C. Bailey and Wesley G. Span- 
genberg, Service Projects AC-l, NO-115, and others, NBS. 

Div. 5-350-M4 

Appendix C 

1. OSRD Contractor's Report, J. Strong, D4-4, p. 13. 
la. Ibid., Ref. 1, Fig. 21. 

2. The Review of Scientific Instruments, I I. V. Hayes, Vol. 8, 
1937, p. 342. 

Physics, E. O. Ilulburt, Vol. 5, 1934, p. 101. 

Journal of the Optical Society of America, Smith and H. V. 
Hayes, Vol. 30, 1940, p. 332. % 

3. Annalen der Physik, Mie, Vol. 25, 1908, p. 377. 

4. II. G. Houghton, W. H. Radford, and A. C. Bemis, MIT, 
Cambridge, Mass. 

5. The Physical Review, L. P. Granath and E. O. Ilulburt, 
Vol. 34, 1929, p. 140. 

6. Astrophysical Journal, A. Adel, Vol. 89, 1939, p. 1; Vol. 
91, 1940, p. 1; Vol. 91, 1940, p. 481. 

The Physical Review, W. M. Elsasser, Vol. 54, 1938, p. 
126. 

Journal of the Franklin Institute, J. Strong, Vol. 232, 
1941, p. 1. 

Papers in Physical Oceanography and Meteorology, F. A. 
Brooks, MIT, Vol. 8, October 1941; Science Abstracts, 
Vol. 45, 1942, p. 107. 

7. Astrophysical Journal, A. Adel, Vol. 96, No. 2, 1942, p. 
239. 

The Physical Review, P. E. Martin and E. F. Barker, 
Vol. 41, 1932, p. 291. 

8. Progress Report OEMsr-60, J. Strong, Dec. 31, 1942. 

The Physical Review, L. Harris, Vol. 45, 1934, p. 635. 

9. Journal of the Franklin Institute, P. Moon, Vol. 219, 1935, 
p. 17. 

10. The Review of Scientific Instruments, Andrews, Bruksch, 
Ziegler, and Blanchard, Vol. 13, 1942, p. 281. 

11. Journal of the Optical Society of America, P. Moon and 
L. R. Steinhardt, Vol. 28, 1938, p. 148. 

12. The Review of Scientific Instruments, H. V. Hayes, Vol. 7, 

1936, p. 202. 

13. Zeitschrift fur Physik, M. Czerny and P. Mollet, Vol. 108, 

1937, p. 85. 



OSRD APPOINTEES 


division 5 

Division Chiefs 


Harold B. Richmond 


Hugh H. Spencer 


Division Members 


Jerome C. Hunsaker 
Alfred L. Loomis 
Pierre Mertz 
Harold B. Richmond 
Louis N. Ridenour 

Consultants 


Joseph C. Boyce 
Oliver E. Buckley 
Hugh L. Dryden 
Lars O. Grondahl 
Frederick L. Hovde 


Frank J. Bingley 
Englehardt A. Eckhardt 
Laurens Hammond 
O. B. Hanson 
Daniel E. Harnett 
Edwin H. Land 

John D. 


George A. Philbrick 
Edward M. Purcell 
William H. Radford 
Harold B. Richmond 
Robert E. Shelby 
Donald B. Sinclair 

Williams 


Section Chiefs 


Joseph C. Boyce 
Oliver E. Buckley 
Hugh L. Dryden 


Lars O. Grondahl 
Jerome C. Hunsaker 
Pierre Mertz 


Section Members 


Alan C. Bemis 
Frank J. Bingley 
Henry Blackstone 
Joseph C. Boyce 
Oliver E. Buckley 
Hugh L. Dryden 
Englehardt A. Eckhardt 
Lars 0. Grondahl 
O. B. Hanson 

Bertram E. 


Daniel E. Harnett 
Jerome C. Hunsaker 
Ernest M. Lyman 
Pierre Mertz 
George A. Philbrick 
Robert E. Shelby 
Donald B. Sinclair 
Hugh H. Spencer 
John D. Strong 
Warren 


Technical Aides 


Roger Cutting 
James W. Fitzwilliam 
Albert F. Murray 


Earl W. Phelan 
Hugh H. Spencer 
Abner J. Wollan 


CONTRACT NUMBERS, CONTRACTORS, AND SUBJECT OF CONTRACTS 


Contract Number 


Name and Address of Contractor 


Subject 


Transfer of funds 

NDCrc-96 

NDCrc-173 

NDCrc-180 

NDCrc-183 

NDCrc-188 

OEMsr-141 

OEMsr-171 

OEMsr-187 

OEMsr-240 

OEMsr-278 

OEMsr-286 

OEMsr-298 

OEMsr-327 

OEMsr-441 

OEMsr-476 

OEMsr-513 

OEMsr-514 

OEMsr-515 

OEMsr-615 

OEMsr-620 

OEMsr-694 

OEMsr-727 

OEMsr-921 

OEMsr-978 


National Bureau of Standards 
Washington, D. C. 

Hazeltine Electronics Corporation 
Washington, D. C. 

Radio Corporation of America 
New York, New York 

The Massachusetts Institute of Technology 
Cambridge, Massachusetts 

Gulf Research and Development Company 
Pittsburgh, Pennsylvania 

The Massachusetts Institute of Technology 
Cambridge, Massachusetts 

Radio Corporation of America 
New York, New York 

Radio Corporation of America 
New York, New York 

Remington Rand, Inc. 

New York, New York 

The Massachusetts Institute of Technology 
Cambridge, Massachusetts 

Purdue University 
Lafayette, Indiana 

Hazeltine Electronics Corporation 
Washington, D. C. 

Radio Corporation of America 
New York, New York 

Douglas Aircraft Company 
Santa Monica, California 

Radio Corporation of America 
New York, New York 

Vidal Corporation 
Camden, New Jersey 

Radio Corporation of America 
New York, New York 

Radio Corporation of America 
New York, New York 

Radio Corporation of America 
New York, New York 

Radio Corporation of America 
New York, New York 

Farnsworth Television and Radio Corporation 
Fort Wayne, Indiana 

Hammond Research Corporation 
Gloucester, Massachusetts 

Dalmo-Victor Company 
San Carlos, California 

Electro-Mechanical Research, Inc. 

Houston, Texas 

Eastman Kodak Company 
Rochester, New York 


Glide bomb 
Television 

Television, gliders, aerial torpedo 

Heat control 

Bombs 

Bombs 

Glide-bomb controls 
Television 

Television apparatus 
Engineering, bombs 
Radio control for model airplanes 
Television 

Television equipment 

Dive bomb 

Television 

Gliders 

Television 

Television 

Television 

Television 

Television equipment 
Radio control link 
Electromechanical coupling devices 
Scanning, stabilization, and bolometers 
Data recording cameras 


369 












CONTRACT NUMBERS, CONTRACTORS, AND SUBJECT OF CONTRACTS ( Continued) 


■= T-— 

Contract Number 

Name and Address of Contractor 

Subject 

OEMsr-1002 

Bendix Aviation Corporation 

Detroit, Michigan 

Amplifier, Roc 

OEMsr-1013 

The Massachusetts Institute of Technology 
Cambridge, Massachusetts 

Stabilization, servomechanism 

OEMsr-1068 

General Mills, Inc. 

Minneapolis, Minnesota 

Organic homing device 

OEMsr-1081 

Union Switch and Signal Company 

Pittsburgh, Pennsylvania 

1,000-lb Azon 

OEMsr-1093 

Farnsworth Television and Radio Corporation 
Fort Wayne, Indiana 

Television camera 

OEMsr-1159 

Philco Corporation 

Philadelphia, Pennsylvania 

Television radio link 

OEMsr-1172 

General Electric Company 

Schenectady, New York 

Television radio link 

OEMsr-1180 

Timm Aircraft Corporation 

Van Nuys, California 

Production design, gliders 

OEMsr-1182 

Fairchild Camera and Instrument Company 
Jamaica, New York 

Scanning system 

OEMsr-1191 

Columbia Broadcasting System 

New York, New York 

Comparison of television cameras 

OEMsr-1195 

Harvey Radio Laboratories, Inc. 

Cambridge, Massachusetts 

Radio link and Razon simulator 

OEMsr-1258 

Remington Rand, Inc. 

New York, New York 

Heat homing bomb, Felix production 

OEMsr-1274 

Norton Company 

Worcester, Massachusetts 

Heat homing bomb, Felix production 

OEMsr-1285 

Union Switch and Signal Company 

Pittsburgh, Pennsylvania 

2,000-lb Azon 

OEMsr-1287 

Fairchild Camera and Instrument Corporation 
Jamaica, New York 

Felix scanning head 

OEMsr-1301 

General Instrument Corporation 

Elizabeth, New Jersey 

Special amplifier, Felix 

OEMsr-1314 

Philco Corporation 

Philadelphia, Pennsylvania 

Razon receiver 

OEMsr-1317 

Polaroid Corporation 

Cambridge, Massachusetts 

Scanning device 

OEMsr-1348 

General Electric Company 

Schenectady, New York 

Thermal elements 

OEMsr-1402 

Remington Rand, Inc. 

New York, New York 

Felix production design 

OEMsr-1415 

Union Switch and Signal Company 

Pittsburgh, Pennsylvania 

1,000-lb Razon 

OEMsr-1445 

Bendix Aviation Corporation 

Detroit, Michigan 

Roc amplifier 

OEMsr-1451 

Remington Rand, Inc. 

New York, New York 

Felix production 

OEMsr-1454 

L. N. Schwien Engineering Company 

Los Angeles, California 

Control and test instruments 

OEMsr-1493 

Specialties Manufacturing Company, Inc. 

Syosset, New York 

Electronic simulator 


370 













SERVICE PROJECT NUMBERS 



The projects listed below were transmitted to the Executive Secretary, 

NDRC, from the War or Navy Department through either the War De¬ 
partment Liaison Officer for NDRC or the Office of Research and Inven¬ 
tions (formerly the Coordinator of Research and Development), Navy 

Department. 

Service 

Project 

Numbers 

Subject 

AC-1 

Army 

Improvement of precision in bombing, bombing through overcast. 

AC-36 

Controlled trajectory bombs. 

AC-41 

Radio control of model aircraft. 

AC-42 

Development of a radar system and equipment for controlling target-seeking bombs. 

AC-51 

Development of radar system and auxiliary equipment for controlling target-seeking bombs. 

OD-98 

Investigation and development of rocket target.* 

SC-26 

Control equipment for glide bombs. 

SC-49 

Moth. 

NA-109 

Navy 

Radio controlled aircraft. 

NA-116 

Two-inch dissector tube. 

NA-162 

Standardization of output requirements of electronic homing devices. 

NA-183 

Sonic control of guided missiles.* 

NA-190 

Mimo (miniature Image Orthicon) tube. 

NA-228 

Investigation of X and S band antenna patterns for guided missile applications. 

NA-238 

Development of a flight table. 

NO-40 

Controlled trajectory bombs. 

NO-115 

Radar homing bomb. 

NO-169 

Guided missiles—Project Bat. 

NO-174 

Bat. 

NO-235 

Development of a homing missile in the microwave radar region. 

NO-257 

Dove. 

NS-132 

Television cameras. 

NS-136 

Counter-measures against guided missiles. 

NS-281 

Use of the Image Orthicon against camouflage and through haze. 


♦These projects were never implemented by the Division. 


371 









































































/ 






• t 























































rr- rr 










INDEX 


The subject indexes of all STR volumes are combined in a master index printed in a separate volume. 
For access to the index volume consult the Army or Navy Agency listed on the reverse of the half-title page. 


Absorption of infrared radiation by 
gases, 350 
Acceleration 

handicap to pursuer in dogfight, 323 
in ballodromic pursuit, 298-300 
in scopodromic pursuit, 289-293 
of line of sight, 278 
Acoustic deadening material, 119 
Aerial photography, infrared radiation 
utilized, 349 
Aerodynamics 

limitations on bomb size and shape, 
264 

of Azon and Razon bomb, 35-38, 213 
of glide bombs, 9-12 
of homing missiles, 236-243 
of photoelectric homing bomb, 
168-169 

of radial surfaces, 260 
supersonic, 283 
Aero-homing missiles 
see Homing missiles 
AGC circuits of television receivers, 103 
AGC system for radar homing control, 
23-25 

Aluminum, sun reflecting targets, 62 
AM receiver at 775 me, 113 
AM transmitter at 775 me, 111-113 
American Junior Aircraft Co.’s radio 
control systems, 134-135 
Amplitude modulation vs. frequency 
modulation, 111-118 
frequency modulation inferior, 111 
picture quality results compared, 114 
transmission at 775 me, 111-113 
transmission at 1200 me, 114 
transmission at 1850 me, 114-117 
A/N-CRW-7 receiver, 139 
Angle of attack 

dirigible high-angle bombs, 262 
Felix bomb, 51 
glide bombs, 12 

influence on ballodrome pursuit, 
300-301 

mathematical analysis, 321-322 
Roc bomb, 73-74, 277 
Angle of inclination of glide bombs, 16 
Animals as guided missile operators, 
234 

Antitank missile X-7; 268 
Armor-piercing bombs, 47, 265 
Atmospheric particles, 349-350 
Attenuation of infrared radiation, 
349-350 

Automatic control systems, servo 
theory, 149-156 


absolute stability criterion, 154-155 
frequency response, 149-152 
system sensitivity determined, 
155-156 

transfer function, 152-155 
Automatic pilots for glide bombs, 
16-17 

Azon, 27-47, 211-217, 224-227 
combat performance, 44-47 
design restriction, 27 
engineering contracts, 224-225, 227 
field tests, 43-44 
lift and maneuverability, 35-38 
production, 44 
sighting and parallax, 38-40 
Azon, components, 40-42, 134-137 
battery power, 225 
camera, 166 
flares, 42 
fuze arming, 42 
radio, 40-42, 134-135 
receivers, 136-137 
servo links, 42 
storage batteries, 227 
Azon, electronic simulators, 211-217 
computer, 213-214 
flare projectors, 214-216 
model 1010 trainer, 214-217 
model 1020 trainer, 217 
oscilloscope trainer, 217 
Azon, roll stabilization, 29-35 
Cartesian control system, 29-31 
cylindrical coordinate control sys¬ 
tem, 29 

gyro developments, 33-35 
roll torques, 31-33 
Azon, trajectories, 28-29 

corrections for excessive curvature, 
29 

curvature, 29 

eclipse method of release, 29 
paths, 28 

Background signal in homing missiles, 
235-236 

Ballodrome approach to bomber flight, 
319-321, 326 

Ballodrome pursuit by fighter planes, 
295-301 

acceleration, 298-300 
angle of attack, 300-301 
gun fixed, 300-301 
gun parallel to path, 295-300 
multiple approach, 326 
range versus azimuth, 298 
time factor, 298-300 


Bat glide bomb 

control system, 149, 156 
general description, 7 
reception difficulties, 23 
target discrimination, 233-234 
Beating effects in radio control systems, 
143 

Beetle target seeker, 196-198 
Bell Telephone Laboratories 
glide bomb development, 224 
send-receive bomb (SRB), 23-25 
Bendix “Beetle”, target seeker, 196-198 
Birds as target seekers, 199-201 
conditioning of birds, 199-200 
organic control systems, 201 
servo links for pigeon-controlled 
bombs, 200 

“Blocking effects” of radio-control 
systems, 143 
Bolometers 

disturbances caused by varying heat 
of target, 68 

electronic operation, 56-57 
sensitivity increased, 53 
sensitivity measurements, 188 
Bolometers, types 

Dove Eye’s thermistor bolometer, 68 
gold bolometers, 53, 187-188 
infrared receivers, 352 
Hammond platinum bolometer, 354 
metallic bolometers, 354-355 
nickel strips, 53 
Bomber flight, analysis 

see Interception and escape of 
bombers 

Bomber maneuvering techniques, 
322-328 

decoy action, 326 
dogfight, 323-324 

formation density and shape 327- 
328 

head-on parry, 324 
lateral attack, 326-327 
mass raid tactics, 326-328 
rearward fire vs. forward fire, 322-323 
slant gun maneuvers, 325 
summary, 325 

veering away from interception, 
322-323 

vertical escapes, 324-325 
Bomber mock interception, 332-334 
practice maneuver patterns, 333- 
334 

range scale, 333 

scale rules, 332 

stretched time scale, 332-333 


373 


374 


INDEX 


summary, 287 
Bombs, controlled 
see Guided missiles 
Bore-sight error, definition, 11 
Bore-sight errors of homing missiles, 
252-255 
causes, 253 
formulas, 254-255 
reduction, 253-254 

Brachydrome pursuit by fighter planes, 
302-310 

approach, 302-303 
bullet density, 308-310 
firing range and lead, 302-304 
lead angle, 308 
multiple approach, 325 
slant brachydrome, 310 
Brightness in television picture, 98 
Broadcasting equipment for prewar tel¬ 
evision, 96 

Bullet deflection, bombers vs. fighters, 
322 

Bullet trajectory drop, 322 

California Institute of Technology, 
wind tunnel tests, 338 
Cameras, motion picture, 158-166 
analytic aid for guided missile 
studies, 158 

bomb-trajectory cameras, 164-166 
glide bomb cameras, 159-161 
GSAP (gunsight aiming point) cam¬ 
eras, 161 

missile-borne, 159-162 
nose cameras for dirigible bombs, 
161-162 

open cameras, 163-164 
phototheodolites, 162-163 
slit cameras, 166 

synchronized ground cameras, 
164-166 

Cameras for television 
see Television cameras 
Carbon dioxide for absorbing infrared 
rays, 350 

Carp apparatus for Roc bomb, 273 
CE-47 quadrant photocell, 174 
Classification of missiles, 231 
Clinoballodromic approach of inter¬ 
cepting fighters, 326 
Clinoscopodromes, 295 
Clinoscopodromic approach to bomber 
flight, 318-319 

Clouds, penetration by infrared radia¬ 
tion, 349-350 
Computers, simulative 
see Simulators 

Contrast in television picture, 97 
Crab sight, Mark 15 bombsight attach¬ 
ment, 29, 38-40 
Cruciform design 


dirigible high-angle bombs, 263 
homing bomb, 62 
CW radio target seeker, 196-198 
Cylindrical bomb surfaces, aerody¬ 
namics, 262 

D-C bridge system, infrared electron¬ 
ics, 355 

Decoy action in mass bomber raids, 
326 

Degrees of freedom 
bomb missiles, 9-10 
free flying missiles, 268 
homing missiles, 236 
Derivative-taking bomb Eye, 67 
Differential radiation of heat, 348 
Dirigible high-angle bombs, 257-265 
see also Azon; Razon 
aerodynamic control limitations, 264 
angles of attack and maneuverabili¬ 
ty, 262 

armor-piercing bombs, 265 
bomb-bay carriage, 257 
control factors, 257 
cruciform fin design, 263 
jet-controlled missiles, 264-265 
lift force, 257-258 
maneuverability, 258, 262 
nose cameras, 161-162 
radio-control system, 134-135 
size limitations, 258, 263 
velocity limitations, 263 
weight restrictions, 259 
yaw oscillations, 258-259 
Dirigible high-angle bombs, roll torque, 

259- 262 

cylindrical shroud surfaces, 262 
impulsive occurence of torques, 261 
magnitude of torques, 261-262 
radial asymmetry of bomb, 260 
reduction of roll, 32-33 
roll control problem, 259-260 
shading effects of bomb’s structure, 

260- 261 

Dirigible high-angle bombs, television 
guidance, 120-127 
dummy missile, test use, 123 
failure of contractors to produce sat¬ 
isfactory bomb, 120 
RCA equipment, drop tests, 121-127 
steering problem, 121, 127 
television picture of terrain, 126 
Dogfight bomber maneuver, 323-324 
acceleration handicap of pursuer, 323 
circular flight tracking, 323 
escape initiative, 324 
Douglas Aircraft Co’s, photoelectric 
target seeker, 172-174 
Dove Eye scanner, 65-68 
derivative-taking Eye, 67 
infrared detection, 65-67 


“lock-on” Eye, 67-68 
optical system, 67 
points at target, 67 
thermistor bolometer, 68 
Drag coefficient 
bomber flight, 318 
equation and definition, 35 
Felix bomb, 51 
glide bombs, 12 
influences sighting, 35 

Electromagnetic controls for automat¬ 
ic pilots, 17 

Electro-Mechanical Research, Inc. 
roll stabilizer, 180-185 
scanners, 185-196 
Electronics, infrared, 355-357 
Elevons 

displacement, 13 

electromagnetically controlled by 
automatic pilot, 17 
glide bombs, 10 
Pelican missile, 146-147 
Equilibrium of bomb missiles, 10 
Equilibrium of glide bombs, 12-13 
Error signal, basis for servomechanism 
operation, 146 

Escape techniques of bombers 
see Interception and escape of 
bombers 

Fairchild Camera and Instrument 
Corp., photoelectric scanner, 
174-180 

Farnsworth Television and Radio 
Corp., television equipment for 
high-angle bombs, 120-121 
Feedback amplifiers, analogous to 
servomechanisms, 146 
Felix bomb, 48-65, 68-72 
dynamics, 51-53 
engineering contracts, 225-226 
equations of motion, 51-52 
future prospects, 70-72 
on-off control, 52-53 
radiation theory, 48-49 
scanning system, 50, 53-56 
simulator, 205-207 
target discrimination, 68 
target survey instrument, 68-70 
Felix bomb, electronics, 56-61 
commutator and integrator, 59-61 
operational principles, 56-58 
phase inverter and clipper, 59 
preamplifier, 58-59 
servomotors and feedback, 61 
transformer and input circuit, 
58-59 

Felix bomb, field tests, 61-65 

aluminum target to reflect sun, 62 
bomb shackle failures, 65 



INDEX 


375 


climate testing of components, 64-65 
initial tests, Eglin Field, 61-62 
moisture detrimental to design, 
64-65 

octogonal empennage bomb, 62-63 
preproduction units tested, 64-65 
target design, 64 

target evaluation instrument, 68-70 
targets contrasting with water, 65 
Felix bomb, scanning system, 53-56 
bolometer, 53 
fast parabola optics, 54 
fogging prevention, 55-56 
mutational character, 54 
optical system, 54-55 
scanning head elements, 55-56 
sensitivity, 54-55 
summary, 50 

Fighter planes’ interception 

see Interception and escape of 
bombers 

Flap angle of glide bombs, 12 
Flap control law, Roc bomb dynamics, 
76-77 

Flares for Azon and Razon, 42 
Flicker in television picture, 98 
FM receiver on 775 me, 114 
FM receiver on 1850 me, 116 
FM transmitter on 775 me, 113 
FM transmitter on 1850 me, 114 
Fog penetration by infrared radiation, 
349-350 

Fogging of window surfaces, prevention, 
55-56 
Free gyros 

comparison with rate gyros, 16-17 
roll orientation determination, 33-35 
roll stabilization, 29-31 
Frequency modulation vs. amplitude 
modulation, 111-118 
FM inferior, 111 
transmission at 775 me, 111-113 
transmission at 1200 me, 114 
transmission at 1850 me, 114-117 
Frequency response, servomechanism, 

149- 152 

amplitude response, 150-151 
analysis problem, 149-150 
frequency-response characteristic, 

150- 152 

phase response, 150-151 
relation to transfer function, 152-153 
sinusoidal and transient character¬ 
istic correlated, 151 
sinusoidal response, 150-152 
transient response, 150-152 
FX-1500, visually guided German 
bomb, 29 

Gas radiometers, 352 
Geometry of interception, 278-280 


German V—1 launching sites, targets 
for heat-homing bombs, 62 
Glide bombs, 7-26 
automatic pilots, 16-17 
cameras, 159-161 

coordination of missile and control, 
7-9 

definition, 7 

engineering contracts, 224 
moving targets and wind, 11-12 
radar-homing control, RHB con¬ 
trol, 17-18 

radar-homing control, SRB control, 
23-25 

radio-control systems, 134 
servo links, 25-26 

target discrimination and tracking, 9 
wing loading, 12 

Glide bombs, aerodynamics, 9-12 
angle of attack, 12 
degrees of freedom of missile, 9-10 
drag coefficient, 12 
elevons, 10 

equilibrium seldom reached, 10 
flap angle, 12 
glide angle, 12 
glide path curve, 10 
lift coefficient, 12 
phugoid period, 10 
summary, 240-241 
Glide bombs, stability, 11-16, 
335-347 

fixed control surfaces, 336-338 

instability, 11 

lateral stability, 14-16 

lateral stability equations, 335-336 

longitudinal stability, 12-14 

motion components, 15 

roll hunting system, 345-347 

roll stabilization system, 343-344 

spiral stability, 15 

Glide bombs, television control, 
118-119 

AAF GB-4 bomb, 119 
acoustic deadening covering trans¬ 
mitters, 119 

airborne receiver and control, 
118-119 

radio control transmitter ground lo¬ 
cated, 118-119 
Robin bomb tests, 118-119 
Gliders, SWOD 
see SWOD gliders 
Gold bolometers, 53, 187-188 
Gold bug computor, 205-207 
GSAP (gunsight aiming point) cam¬ 
eras, 161 

Guided missile attack phases, 93 
Guided missile classifications, 266 
Guided missiles 

Azon and Razon, 27-47, 211-227 


dirigible high-angle bombs, 257-265 
Felix bomb, 48-72 
glide bombs, 7-26 
homing missiles, 231-256 
photoelectric homing bombs, 
167-171 

Roc bomb, 73-93, 266-283 
Gulf Research and Development Co. 
photoelectric homing bomb, 167-171 
radio control system, 134-135 
target seeking test table, 203-205 
Gyros 

bias gyro, 340 
free versus rate, 16-17 
roll orientation use, 33-35 
roll stabilization use, 29-31 
SWOD gliders, 346-347 
turn gyro, 339-340 

Hammond platinum bolometer, 354 
Hammond radio-control system, 
140-143 

beating effects, transmission dis¬ 
turbance, 143 

blocking effects, transmission dis¬ 
turbance, 143 

construction and operation, 140-143 
shocking effects, transmission dis¬ 
turbance, 143 

superaudible modulation use, 140 
volume control, 143 
Head-on parry combat maneuver, 324 
Heat detection devices, infrared, 
351-357 

background radiation, 352-353 
electronics, 355-357 
optics, 355 

receiver operation, 353-354 
Heat radiation, 48-49 

see also Infrared radiation, homing 
missile applications 
absorption as function of wave¬ 
length, 48 

heat measurement, 48-49 
source of radiation, 48 
water vapor absorption of radiation, 
48 

Heat-homing scanner, 65-68 
Homing missiles, 231-256 
bore-sight errors, 252-255 
classification, 231 
launching problems, 256 
pursuit curve error equations, 
254-255 

pursuit curves, 250-255 
radio transmission, 170 
range, 255 
sensitivity, 254-255 
strength problems, 255 
Homing missiles, aerodynamics, 
236-243 




376 


INDEX 


see also Homing missiles, stability 
“complete equilibrium”, 228 
degrees of freedom, 236 
equilibrium, 237-238 
force components, 236 
lateral acceleration and force, 
238-239 

radii of curvature, 239 
rectilinear flying speed, 239 
reference axes, 237 
trimming, compensation for unbal¬ 
ance, 237-238 

Homing missiles, controls, 243-245 
error signals, 244-245 
intelligence coordination with con¬ 
trols, 243-244 

interdependence of controls, 243 
reduction of interaction between con¬ 
trols, 245 

roll, effect on error signals, 244-245 
summary, 237 

Homing missiles, stability, 245-250 
see also Homing missiles, aerody¬ 
namics 

antihunt devices, 248-249 
disturbed motions, 243 
mathematical analysis, 249-250 
models used for analysis, 250 
oscillations of missile, 245-247 
spinning motion, 243 
time lag compensating devices, 
248-249 

time lag in hunting, 247-248 
time to come on course, 249 
wind gusts as instability factor, 249 
Homing missiles, target discrimination, 
232-236 

animals as intelligence devices, 234 
background signal, 235-236 
directional sensitivity, 233 
field of view, 234-235 
limitations, 232 
range selection, 234 
servomechanism action, 235 
signal intensity fluctuations, 236 
signal strength selection, 233 
vehicle motion, 235 
Homing pursuit 

see Scopodrome pursuit 
Hytemco wire, 58 

Iconoscope 

compared with image orthicon tube, 
108-109 

operation principles, 98 
sensitivity, 108-109 
Image dissector, television camera tube, 
98 

Image orthicon tube, 100-101, 108-111 
comparison with iconoscope, 108-109 
comparison with vericon tube, 110 


electronic operations, 100-101 
infrared sensitivity, 110-111 
mounting of glass target, 110-111 
multiplying electrode, 111 
photoelectric operation, 98-99 
sensitivity, 108-109 
signal-to-noise ratio, 111 
target, 101 
target screen, 111 

Incidence control in Roc bomb, 269-270 
Infrared heat detection devices, 351- 
357 

automatic control, homing missiles, 
352 

background radiation, 352-353 
electronics, 355-357 
methods of operating receivers, 353- 
354 

optics, 355 

Infrared radiation, attenuation, 349- 
350 

scattering, 349 

solid particles in atmosphere, 349 
water droplets in atmosphere, 349- 
350 

Infrared radiation, homing missile ap¬ 
plications, 348-357 
absorption bands due to gases, 350 
bomb control possibility, 61-62 
emission of infrared radiation, 351 
energy loss due to suspended part¬ 
icles, 349-350 
far infrared, 349 
heat-detection devices, 351-357 
near infrared, 349 
radiation theory, 348-349 
Infrared radiation absorption due to 
gases, 350 

Infrared receivers, 351-354 
bolometers, 352 
gas radiometers, 352 
metallic bolometers, 354-355 
operating methods, 353-354 
relative merits of receivers, 351 
thermocouples, 351-352 
Interception and escape of bombers, 
285-334 

ballodrome pursuit, 295-301 
brachydromic passage, 302-304 
collision course, 301 
combat maneuvering techniques, 286, 
322-325 

immunization of personnel to accel¬ 
eration stress, 315 
interception limitations, 287-288 
interceptor passage without banking, 
301-302 

mass raid tactics, 287, 320-328 
mock interception to scale, 287, 332- 
334 

multiple interception, 286, 325-326 


nomenclature, 285 
scopodrome pursuit, 288-295 
summary, 286-287 
technique theory, 310-315 
tomodrome pursuit, 301-302 
Interception and escape of bombers, 
mathematics, 315-322 
angle of attack, 321-322 
ballodromic approach, 319-321 
clinoscopodromic approach, 318-319 
deceleration due to inertia load, 318 
least load factor peak, 317 
scopodromic curve in pursued’s co¬ 
ordinates, 315-316 
time of scopodromic approach, 317- 
318 

Interception by guided missiles, 277- 
280 

air density and air speed effects, 280 
control system’s adjustment to aero¬ 
dynamic action, 279-280 
geometry of interception, 278-279 
lift coefficient, 280 
maneuvers by remote guidance, 278 
Interceptor fire, probability of success, 
328-332 

factors determining success, 328 
probability calculations, limitations, 
328 

standard conditions assumed, 328- 
331 

summary, 287, 331 

Jag (just another gadget), bombsight 
correcting device, 40 
Jet-propelled projectiles, 264-265 

K-24 camera, 164-166 

Lanchester’s work 

conditions of equilibrium in flight, 
12-13 

phugoid, 9-10 
Lead angle 

brachydrome pursuit, 308 
guided missile attacks, 279 
Lift 

Azon and Razon, 35-38 
control, 268-269 
control speed, 270-271 
dirigible bombs, 257-258 
German control systems, 271 
glide bombs, 10 
maximum lift demands, 270 
Roc bomb, 76-77, 268-271 
Lift coefficient 

dirigible bombs, 258 
equation and definition, 35 
Felix, 51 
glide bombs, 12 

guided missile interception, 280 



INDEX 


377 


Line of sight, guided missiles, 278 
Load factors 

effect on interception success, 329 
scopodromic pursuit, 289-293 
Lock-on bomb Eye, 67 
Looping, bomber maneuver, 324 

Maneuvering techniques of bombers, 
322-325 

dogfight, 323-324 
head-on parry, 324 
maneuver patterns, 333-334 
slant gun maneuvers, 325 
summary, 286, 325 
veering away from interceptor, 322 
vertical escapes, 324-325 
Mass raid bomber tactics, 326-328 
decoy action, 326 
formation density, 327-328 
formation shape, 327 
lateral attack, 326-327 
summary, 287 

Massachusetts Institute of Technology 
see MIT 

Mimo, Roc bomb television 

see Roc bomb, television control 

MIT 

Harvey receiver, 137-139 
RHB development, 22-23 
Servomechanism Lab., 146 
Model range for guidance technique 
tests, 282 

Modulation, frequency vs. amplitude, 
111-118 

Moisture, effect on heat-homing bombs, 
64-65 

Motion picture cameras 

see Cameras, motion picture 
Motion pictures used in photoelectric 
homing tests, 85 

Multipath television transmission prob¬ 
lems, 101-102 

Navigation with television, 94 
Nickel strip bolometers, 53 
Nickle’s transient analyzer, 202 
Norden bombsight, Crab modification, 
29, 38-40 

Nose cameras for high-angle dirigible 
bombs, 161-162 

Nyquist diagram, feedback amplifier, 
154 

Ocular vision from aircraft, 272 
Optics, infrared, 355 
Orthicon tube 

see Image orthicon tube 
Oscillation in yaw, gliders, 340-342 
damped oscillation, 340 
damped vs. divergent oscillation, 
342-343 


periodicity, 340 
predicted damping, 340 
sideslip velocity effects, 342 
summary, 337 

Oscillation period, glide bombs, 15 
Ozone, cause of infrared absorption, 
350 

Parallax between target and missile, 
272 

Parallax correction, bombsights, 38-40 
Pelican bomb, 146-149 
aerodynamic control, 146 
cameras, 160 
elevon motion, 147 
flight tests, 148 
gyros, 146 

reception difficulties, 23 
servo system test table, 148, 210-211 
simulator, 207-210 
target discrimination, 233-234 
Photoelectric homing bombs, 167-171 
see also Roc bomb, photoelectric 
scanning 

aerodynamics, 168 
phototube, 170 

radio transmission from bomb, 170- 
171 

roll stabilization, 167 
scanning system, 168-169 
tests, 170 

Photoreconnaissance data, 70 
Phototheodolites, 162-163 
Phugoid, aerodynamic curve, 10 
Pigeons as target seekers, 199-201 
conditioning of birds, 199-200 
field of view in bombs, 200 
organic control systems, 201 
“reserve of impulses” in bird, 199 
servo links for pigeon-controlled 
bombs, 200 

Polaroid Corporation’s Dove Eye, 65- 
67 

Probability, interceptor fire effective¬ 
ness, 328-332 

assumption of standard conditions, 
328-331 

factors determining success, 328 
limitations to probability calcula¬ 
tions, 328 
summary, 287, 331 
Proportional control link, 144-145 
audio tones, 144 
discriminator, 144 
servomotor difficulties, 145 
unnecessary for Azon, 136 
Pulse method, infrared receiver opera¬ 
tion, 353-354 

Purdue University’s proportional con¬ 
trol link studies, 144-145 
Pursuit curve computers, 252 


Pursuit curves 
for homing missiles, 250-255 
scopodromic, 315-319 
tomodromic, 301-302 
Pyrotechnic flares for Azon and Razon, 
42 

Quadrant photocell target seeker, 172- 
174 

Radar control for Roc bomb, 74-75, 
271-272 

Radar echoes returned from targets, 
19-20 

Radar-homing bomb (RHB), 17-23 
antenna system, 22 
radar illumination of target, 17-19 
range, 22 
receiver, 20-22 
self-synchronous RHB, 23 
target contrast, 19-20 
threshold sensitivity, 22 
tracking circuit, 21 

Radial surfaces’ aerodynamic features, 
260 

Radiation loss in atmosphere, 349-350 
scattering, 349 
solid particles, 349 
water droplets, 349-350 
Radiation sources, target vs. back¬ 
ground, 352-353 
Radio target seeker, 196-198 
Radio transmission from homing bomb, 
170 

Radio-control systems, 134-145 
see also Azon; Razon; Roc bomb 
glide bomb systems, 134 
Hammond control system, 140-143 
high-angle bomb systems, 134-135 
problems of developmental program, 
134 

proportional control, 144-145 
secure systems, 140-143 
transmission disturbances, 143 
Radiosonde oscilloscope, 170-171 
Rate gyros 

damping of roll oscillations, 33-35 
roll stabilization use, 29-31 
versus free gyros, 16-17 
Razon, 27-47, 211-217, 224-227 
design restriction, 27 
engineering contracts, 224-225, 227 
field tests, 43-44 
lift and maneuverability, 35-38 
motion equations, 213 
production, 44 
sighting and parallax, 38-40 
Tarzon, bomb with Razon controls, 
47 

use against armor of capital ships, 47 
Razon, components, 40-42, 134-139 



378 


INDEX 


battery power, 225 
camera, 166 
flares, 42 
fuze arming, 42 
radio, 40-42, 134-135 
receivers, 137-139 
servo links, 42 
storage batteries, 227 
Razon, electronic simulators, 211-217 
computer, 213-214 
flare projectors, 214-216 
model 1010 trainer, 214-217 
model 1020 trainer, 217 
motion equations, 213 
oscilloscope trainer, 217 
Razon, roll stabilization, 29-35 
Cartesion control system, 29-31 
cylindrical coordinate control system, 
29 

gyro developments, 33-35 
roll torques, 31-33 
Razon, trajectories, 28-29 

corrections for excessive curvature, 
29 

curvature, 29 

eclipse method of release, 29 
paths, 28 

RCA, television equipment for diri¬ 
gible high-angle bombs, 121-127 
Remington Rand Co., vericon tube 
production, 110 
Research recommendations 

bombsight for evasive action, 88 
engineering staff for scientific proj¬ 
ects, 223-224 

Felix bomb on-off control, 52-53 
heat-homing missiles, 70-72 
Roc bomb, 283 
supersonic aerodynamics, 283 
transient aerodynamic responses, 
guided missiles, 90 

Resolution of television picture, 96-97, 
104-107 

RHB (radar-homing bomb), 17-23 
antenna system, 20 
radar illumination of target, 17-19 
range, 22 
receiver, 20-22 
self-synchronous RHB, 23 
target contrast, 19-20 
threshold sensitivity, 22 
tracking circuit, 21 
Robin bomb 

radio control system, 134 
structure, 7 

television control tests, 118-119 
Roc bomb, 73-93, 266-283 
antenna, 128-130 
bankless turn, 73-74 
cross-wing structure, 79 
engineering program, 226 


lift, 77 

maneuverability, 128 
nose cameras, 162 
phases of an attack, 93 
research recommendations, 283 
roll stabilization, 173 
zero angle of attack, 73-74 

Roc bomb, guidance, 73-76, 271-273 
see also Roc bomb, television control 
Carp apparatus, 273 
guider station control, 272 
model range for testing, 282 
radar beam rider, 271-272 
radar control, 73-75 
remote guidance, 272 
Skramstad’s control regime, 76 
summary, 79-82 

television and radio control, 75-76 
television control chosen instead of 
radar, 73 

visual guidance, 88 

Roc bomb, photoelectric scanning, 82- 
85, 174-180 

AGC characteristics, 177-178 
aperture of optical system, 177 
homing tests, 82-85 
nonoscillatory characteristics, 82-83 
operational principles, 174-176 
optics, 176-177 
preamplifier, 177-179 
summary, 79-82 
tests, 179-180 

Roc bomb, simulator, 218-222 
aerodynamic equations, 219-220 
computer, 221-222 
input-output relationships, 220 
presentation, 222 

time integration and differentiation, 
220-221 

Roc bomb, television control, 127-133, 
273-283 

aiming equipment, 274-275 
camera, 128, 275 

compensation for angle of attack os¬ 
cillations, 277 

compensation for angle of trim, 277 
computer within aircraft, 280 
gravity computation on missile, 281 
interception maneuvers, 277-280 
‘‘miss” defined, 276 
model range for testing, 282 
remote control steering, 273-274 
target motion compensation, 275-276 
tests, 130-133, 281-283 
two-man control, 274 
wind compensating devices, 275-276 

Roc bomb, wing system, 266-271 
attachment to bomb, 266-267 
incidence control, 269-270 
lift control, 268-271 
maneuverability demands, 270 


roll stabilization, 267-268 
trim stability, 269 
Roc 00-1000 bomb, 85-93, 267 
control system, 87-88 
dimensions and performance con¬ 
stants, 85 

structural design, 85-87 
television guidance tests, 88-93 
visual guidance tests, 88 
Rocket-type jet bombs, 264-265 
Roll 

damping, 15 

effect on error signals, 244-245 
in SWOD gliders, 339 
of bomb missiles, 10 
of photoelectric homing bombs, 167 
of Roc bomb, 173, 267-268 
Roll stabilization in Azon and Razon, 
29-35 

Cartesian control system, 29-31 
cylindrical coordinate control sys¬ 
tem, 29 

gyro developments, 33-35 
Roll stabilization in glide bombs, 343- 
347 

motion in roll, 343-344 
Roc bomb, 73, 267-268 
roll hunt period, 344 
time lag in servo response, 343-344 
Roll stabilization without free gyro, 
180-185 

initial acceleration, 182 
initial velocity, 182 
storage circuit for aileron reversal, 
184 

summary, 185 
tests, 184 
theory, 180-183 
time lags, 183 
Roll torque, 259-262 

effect of bomb’s radial asymmetry, 
260 

effect of cylindrical shroud surfaces, 
262 

impulsive occurrence of torques, 261 
in Azon and Razon bombs, 31-33 
induced roll torques, 31-32 
magnitude, 261-262 
reduction, 32-33 
roll control problem, 259-260 
shading effect of bomb’s structure, 
260-261 

wind tunnel studies, 31-32 

S-3 scanner, 185-188 
S-4 and S-5 scanners, 188 
S-6 scanner, 190 
S-7 scanner, 190-192 
S-8 scanner, 193-196 
amplifier circuits, 194 
construction, 193-194 



INDEX 


379 


synchronous rectifier, 194-196 
Scanners 

Dove Eye, 65-68 
for Felix bomb, 53-56 
Scanners, harmonic, 188-190 
choice of second harmonic, 190 
higher harmonics, 189 
response curves, 188-189 
Scanners, photoelectric, 174-180 
operational principles, 174-176 
optics, 176-177 
preamplifier, 177-179 
Scanners, to-and-fro, 185-196 
bolometers, 187-188 
harmonic scanning, 188-190 
horizon problem, 185 
output, 179 
S-3 scanner, 185-186 
S-4 scanner, 188 
S-5 scanner, 188 
S-6 scanner, 190 
S-7 scanner, 190-192 
S-8 scanner, 193-196 
sensitivity measurements, 188 
theory, 185-187 

Schmidt optics, applied to image orthi- 
con tube, 109-110 
Schwien gyro unit, 281 
Scopodrome pursuit, 288-295 
accelerations, 289-293 
centripetal acceleration, 317 
clinoscopjodromes, 295 
horizontal plane scopodrome, 289 
interceptor’s speed constant, 288 
load factor limitations, 289-293 
pitched gun pursuit, 295 
range versus azimuth, 288-289 
speed loss, 294 

time elapsed before interception, 
292-294 

vertical and oblique scopodromes, 
294 

Scopodromic pursuit curve, 315-319 
clinoscopodromic approach, 318-319 
equation in pursued’s polar coordi¬ 
nates, 315-316 
lead load factor peak, 317 
peak load factor, 316 
time of approach, 317-318 
Sea surface’s effect on background rad¬ 
iation, 235 

Security of radio-control systems, 140- 
143 

Self-propelled missiles, 264-265 
Send-receive bomb (SRB), 23-25 
Servo links, 25-26 

Azon and Razon use, 42 
continuously running electric motor, 
25 

definition, 9 
design problem, 25 


for pigeon-controlled bomb, 200 
linkage, 25-26 
Servo theory, 149-156 

absolute stability criterion, 154-155 
frequency response, automatic con¬ 
trol system, 149-152 
system sensitivity determination, 
155-156 

transfer function, automatic control 
system, 152-153 
transfer function locus, 153-154 
transfer loci, 155 
Servomechanisms, 146-156 

Bat missile control system, 149 
definition, 9, 146 
glide bombs, 11 

MIT Servomechanisms Lab., 146 
Pelican missile control system, 146- 
149 

target discrimination role, 235 
“Shocking effects” disturbing operation 
of radio-control systems, 143 
Shroud surfaces, cylindrical, 262 
Silver chloride, used to prevent win¬ 
dow fogging, 55-56 
Simulators, 202-222 

Azon and Razon electronic simula¬ 
tors, 211-217 
Gold bug, 205-207 
Pelican servo test table, 210-211 
Pelican simulator, 207-210 
photoelectric bomb test table, 203- 
205 

Roc bomb simulator, 218-222 
simulative analysis, applications, 202 
-203 

television guidance representation, 
282 

Sinusoidal frequency response, auto¬ 
matic control systems, 150-152 
Skramstad’s law, control-flap position, 
76-77 

Smoke particles, penetration by infra¬ 
red radiation, 349 

“Snag dodge,” bomber maneuver, 324 
Spectral response of television cameras, 
100 

Spinning motion of missiles, 243 
SRB (send-receive bomb), 23-25 
Steady-state data, 12 
Steady-state method, infrared receiver 
operation, 353-354 

Steering problem for remote control 
missiles, 273-274 

Stefan-Boltzmann law, wavelength ra¬ 
diation, 48 
Suicide missiles, 198 
Sun’s reflection for heat-homing bomb 
attraction, 62 

Superaudible modulation, radio-con¬ 
trol system use, 140 



Supersonic aerodynamics, 283 
Suspended particles in atmosphere, at¬ 
tenuate radiation, 349-350 
SWOD gliders, 338-343 
gyro and servo system, 339-340 
Mark 7 and 9; 339-343 
Mark 12 and 13; 338-339 
oscillation in yaw, 340 
roll axis, 339 

sideslip velocity, 342-343 
stability, 338-339 

time lag in homing intelligence, 340- 
342 

Systems engineering defined, 7-9 

Target contrast, radar, 19-20 
Target discrimination 
Felix bomb, 68 
glide bombs, 9, 233-234 
heat-homing bombs, 68 
infrared radiators, 352-353 
Target discrimination by homing mis¬ 
siles, 232-236 

animals as intelligence devices, 234 
background signal, 235-236 
directional sensitivity, 233 
field of view, 234-235 
limitations, 232 
range selection, 234 
servomechanism action, 235 
signal intensity fluctuations, 236 
signal strength selection, 233 
vehicle motion, 235 

Target discrimination by pigeons, 199- 
201 

organic control systems, 201 
servo links for pigeon-controlled 
bombs, 200 

training program for birds, 199-200 
Target seeking devices, 167-201 

based on infrared radiation, 348-357 
Beetle, radio target seeker, 196-198 
organic methods, 198-201 
photoelectric homing bomb, 167-171 
photoelectric scanner, 174-180 
quadrant photocell target seeker, 
172-174 

to-and-fro scanners, 185-196 
Targets for Felix bomb testing, 62-65 
aluminum target to reflect sun, 62 
design of target, 64 
preparation, 62 

steel plates heated by oil burners, 
62-64 

target evaluation instrument, 68-70 
targets contrasting with water, 65 
Tarzon, deep penetration bomb, 47 
Television bomb guidance, 88-93 
see also Dirigible high-angle bombs, 
television guidance; Roc bomb, 
television control 



380 


INDEX 


AAF GB-4 bomb, 119 
aiming device, 88-89 
complexity of guidance problem, 89- 
90 

general navigation not feasible, 94 
glide bombs, 118-119 
guided missile problem, 100 
phases of an attack, 93 
picture quality of an operation, 94 
picture requirements for accurate 
steering, 94 

radio link requirements, 94-96 
radio-control transmitter, 118 
stages of a maneuver, 94 
Television bomb guidance, equipment, 
96-104 

see also Television cameras 
AGC circuits, 103 
“clamp” circuit, 103-104 
equipment incorporating improve¬ 
ments, 124-127 
head-end tuner, 103 
iconoscope lamp, 103 
image orthicon tube, 101 
noise peaks clipped, 103 
optical system aperture, 100-101 
percentage modulation meter, 102 
prewar equipment, 96-99 
stabilized oscillators, 103 
video amplifier, 103 
Television cameras, 88-111 

acoustic deadening material, camera 
cover, 110 

aperture of optical system, 100-101 
engineering principles applied to 
equipment design, 100 
equipments listed, 104 
iconoscope, prewar, 98 
iconoscope and image orthicon tube 
compared, 108-109 
image orthicon tube, 100-101, 109- 
111 

infrared sensitivity improved, 110 
Mimo, camera for Roc bomb, 127- 
133, 275 

prewar cameras, 98-99 
quantitative evaluation of contrast, 
100 

resolution comparison, 105-108 
Schmidt optics applications, 109-110 


sensitivity comparison, 105-108 
sensitivity improvements, 108-111 
spectral response measured, 100 
storage tubes, prewar, 98-99 
vericon tube, 110 

vulnerability of equipment to audio 
vibrations, 105-107 
Television picture, transmission qual¬ 
ity, 96-98 
brightness, 98 
contrast, 97 
flicker, 98 

overall resolution, 96-97 
synchronization, 96 
Television radio links, 101-102 

frequency modulation versus ampli¬ 
tude modulation, 102 
multipath transmission problems, 
101-102 

transmission at very high frequencies, 
101 

Television transmission at 775 me, 
111-114 

AM receiver and transmitter, 111- 

113 

flight tests, 114 

FM receiver and transmitter, 114 
Television transmission at 1200 me, 

114 

Television transmission at 1850 me, 
114-117 

frequency modulation chosen, 114 
miniature transmission device, 118 
receiver, 116 
test results, 117-118 
transmitter, 114 

Thermal sensitivity of bolometers, 53 
Thermistor bolometer, Dove Eye use, 
68 

Thermocouples, 351-352 
Thermopiles, 53, 353 
‘‘Time derivative operator, ’ ’ Roc bombs, 

77 

Tomodromic pursuit course, 301-302 
Tracking, glide bombs, 9 
Tracking in circular flight, 323 
Trajectories, Azon and Razon, 28-29 
corrections for excessive curvature, 
29 

curvature, 29 


eclipse method of release, 28 
paths, 28 

Trajectory slope of guided missile, 266 
Transfer function, automatic control 
system, 152-153 

Transfer function locus, servomech¬ 
anisms, 153-154 

Transfer loci, servomechanism, 155 
Transient frequency response, auto¬ 
matic control systems, 150-152 
“Trial angle,” Crab sight, 38-40 
Trim in homing missiles, 237-238 
Trim stability, 269 

Union Switch and Signal Company, 
Azon and Razon development, 
224-225 

VB dirigible bombs, 257-265 
see also Dirigible high-angle bombs 
cruciform structure, 263 
oscillation period, 259 
radii of curvature, 257 
size limitations, 258 
VB-13, penetration povrer, 265 
VB-13, size, 263 

Velocity limitations on bombs, 263 
Vericon, orthicon type tube, 110 
Video gain, television picture, 97 
Video signals 

amplitude modulation not feasible, 
102 

prewar television receivers, 99 
spurious frequency modulation, 101— 
102 

Water droplets in atmosphere, 349- 
350 

Water vapor, cause of infrared absorp¬ 
tion, 350 

Water vapor, cause of radiation absorp¬ 
tion, 48 

X-7 German antitank missile, 268 

Yaw oscillation 

dirigible bombs, 258-259 
gliders, 340-342 
homing bombs, 205 

Zero angle of attack, Roc bomb, 73-74 





















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